Turbine engine

ABSTRACT

A gas turbine engine for an aircraft includes an engine core including a turbine, compressor, and core shaft connecting turbine to compressor; a fan located upstream of the engine core and including a plurality of fan blades each having a leading and trailing edge. The turbine includes a lowest pressure turbine stage having a row of rotor blades, each rotor blades extending radially and having a leading and trailing edge. The engine has a fan tip axis that joins a radially outer tip of the leading edge of a fan blade and the radially outer tip of the trailing edge of a rotor blade of the lowest pressure stage. The fan tip axis lies in a longitudinal plane which contains a centreline of engine. A fan axis angle is defined as the angle between fan tip axis and centreline, and is in a range between 10 and 20 degrees.

CROSS-REFERENCE TO RELATED APPLICATION(S)

This application is a continuation of U.S. application Ser. No.16/825,361 filed on Mar. 20, 2020, which is a continuation of A.S.application Ser. No. 16/671,215 filed Nov. 1, 2019, which is based onand claims priority under 35 U.S.C. 119 from British Patent ApplicationNo. 1820924.7 filed on Dec. 21, 2018, and British Patent Application No.1820918.9 filed on Dec. 21, 2018. The contents of the above applicationsare incorporated herein by reference.

BACKGROUND

The present disclosure relates to a gas turbine engine for an aircraft,and more specifically to a gas turbine engine with specified relativecomponent dimensions.

The skilled person would appreciate that simply scaling up components ofa known engine type may not provide a corresponding scaling ofpower/thrust and/or efficiency, and may introduce problems such asincreased drag or difficulty of installation. Reconsideration of engineparameters may therefore be appropriate.

For example, the skilled person would appreciate that, if the overallsize of a gas turbine engine is increased, one problem that may need tobe addressed is how to reduce the overall drag produced by acorrespondingly larger nacelle of the larger engine when it is in use.If the components of the engine are scaled proportionally—simply scalingup a known engine type—the increased drag may negatively affect theperformance of the aircraft on which the engine is mounted. Additionallyor alternatively, the engine may not fit to be mounted beneath the wingof the aircraft unless dimensions are adjusted.

As used herein, a range “from value X to value Y” or “between value Xand value Y”, or the likes, denotes an inclusive range; including thebounding values of X and Y. As used herein, the term “axial plane”denotes a plane extending along the length of an engine, parallel to andcontaining an axial centreline of the engine, and the term “radialplane” denotes a plane extending perpendicular to the axial centrelineof the engine, so including all radial lines at the axial position ofthe radial plane. Axial planes may also be referred to as longitudinalplanes, as they extend along the length of the engine. A radial distanceor an axial distance is therefore a distance in a radial or axial plane,respectively.

According to one aspect there is provided a gas turbine engine for anaircraft, the engine comprising an engine core having a core length andcomprising a turbine, a compressor, and a core shaft connecting theturbine to the compressor, the turbine comprising a lowest pressurerotor stage, the turbine having a turbine diameter at the lowestpressure rotor stage; and a fan located upstream of the engine core, thefan comprising a plurality of fan blades extending from a hub, the huband fan blades together defining a fan face having a fan face area and afan tip radius. An engine area ratio of:

$\frac{{the}\mspace{14mu}{fan}\mspace{14mu}{face}\mspace{14mu}{area}}{\begin{matrix}{{the}\mspace{14mu}{turbine}\mspace{14mu}{{diameter}\mspace{14mu}\left( {{at}\mspace{14mu}{the}{\mspace{11mu}\;}{lowest}\mspace{14mu}{pressure}\mspace{14mu}{rotor}\mspace{14mu}{stage}} \right)} \times} \\{{the}\mspace{14mu}{core}\mspace{14mu}{length}}\end{matrix}}$

is in the range from 1.7 to 3.

The present aspect relates to a gas turbine engine with specifiedrelative fan face and engine sizes. The skilled person will appreciatethat a larger fan may provide improved propulsive efficiency. Theskilled person will appreciate that a relatively small turbine diameter,compared to the fan size, may improve ease of installation. The skilledperson will appreciate that a relatively short core length, and/or arelatively narrow core diameter, compared to the fan size, may benefitclose coupled installation.

The skilled person will appreciate that turbine diameter×core length mayprovide an effective engine area in an axial plane, and that reducingthis area may benefit close-coupled installation. In particular, the fanmay be mounted closer to the wing when the engine area is smaller(mounted further back and further up than otherwise), so reducing themoment applied to the wing by the mass of the engine.

The skilled person will appreciate that equivalent units should beselected for the fan face area, turbine diameter and core length—e.g. ifthe area is given in m², the lengths should both be provided in metres.

The engine area ratio may be higher than that of known aircraft gasturbine engines.

The engine area ratio may be in the range from 1.7 to 3.0, andoptionally from 1.70 to 3.00. The engine area ratio may be in the rangefrom 1.8 to 3, or 1.9 to 3 (or optionally to 3.0).

The engine area ratio may be in the range from 2 to 3. The engine arearatio may be in the range from 2.1 to 2.7.

The fan tip radius may be measured between a centreline of the engineand an outermost tip of each fan blade at its leading edge—this mayequivalently be described as the fan tip radius being defined as theradial distance between a centreline of the engine and an outermost tipof each fan blade at its leading edge. The fan face area may be equal toπ multiplied by the square of the fan tip radius.

The fan tip radius, measured between a centreline of the engine and anoutermost tip of each fan blade at its leading edge, may be in the rangefrom 95 cm to 200 cm, for example in the range from 110 cm to 150 cm, oralternatively in the range from 155 cm to 200 cm. The fan tip radius maybe greater than any of: 110 cm, 115 cm, 120 cm, 125 cm, 130 cm, 135 cm,140 cm, 145 cm, 150 cm, 155 cm, 160 cm, 165 cm, 170 cm, 175 cm, 180 cm,185 cm, 190 cm or 195 cm. The fan tip radius may be around 110 cm, 115cm, 120 cm, 125 cm, 130 cm, 135 cm, 140 cm, 145 cm, 150 cm, 155 cm, 160cm, 165 cm, 170 cm, 175 cm, 180 cm, 185 cm, 190 cm or 195 cm. The fantip radius may be greater than 160 cm.

The fan tip radius may be in the range from 95 cm to 150 cm, optionallyin the range from 110 cm to 150 cm, optionally in the range of from 110cm to 145 cm, and further optionally in the range from 120 cm to 140 cm.

The fan tip radius may be in the range from 155 cm to 200 cm, optionallyin the range from 160 cm to 200 cm, and further optionally in the rangefrom 165 cm to 190 cm.

Optionally, for example for an engine with a fan tip radius in the rangefrom 110 cm to 150 cm, the engine area ratio may be in the range from1.7 to 3, optionally 1.7 to 2.7, optionally from 2.1 to 2.7, and furtheroptionally from 2.2 to 3.

Optionally, for example for an engine with a fan tip radius in the rangefrom 155 cm to 200 cm, the engine area ratio may be in the range from 2to 3, optionally 2.2 to 3, optionally 2.3 to 2.6, and optionally from2.5 to 2.6.

The turbine diameter at the lowest pressure rotor stage may be measuredat the axial location of blade tip trailing edges of rotor blades of thelowest pressure rotor stage. In embodiments in which the rotor (of thelowest pressure rotor stage) is shrouded, the turbine diameter of theturbine at the lowest pressure rotor stage may be measured to theunderside of the shroud. In embodiments in which the rotor (of thelowest pressure rotor stage) is unshrouded, the turbine diameter of theturbine at the lowest pressure rotor stage may be measured to the bladetips of the rotor. The lowest pressure rotor stage may be the mostaxially rearward (or most downstream) rotor stage.

The turbine diameter at the lowest pressure rotor stage may be in therange from 70 cm to 170 cm. Optionally, for example for an engine with afan tip radius in the range from 110 cm to 140 cm, the turbine diameterat the lowest pressure rotor stage may be in the range from 70 cm to 120cm, for example 80 cm to 115 cm. Optionally, for example for an enginewith a fan tip radius in the range from 155 cm to 200 cm, the turbinediameter at the lowest pressure rotor stage may be in the range from 120cm to 170 cm, for example 130 cm to 160 cm.

The ratio of the fan tip radius to the turbine diameter at the lowestpressure rotor stage

$\left( \frac{{fan}\mspace{14mu}{tip}\mspace{14mu}{radius}}{{turbine}\mspace{14mu}{diameter}} \right)$may be in the range of 0.8 to 2.1. Optionally, for example for an enginewith a fan tip radius in the range from 95 cm to 150 cm, the ratio ofthe fan tip radius to the turbine diameter may be in the range 0.8 to2.1. Optionally, for example for an engine with a fan tip radius in therange from 155 cm to 200 cm, the ratio of the fan tip radius to theturbine diameter may be in the range of 0.9 to 1.7.

The core length may be defined as the axial distance between a forwardregion of the compressor and a rearward region of the turbine. The corelength may be measured along a centreline of the engine from a meanradius point of the first stage of the compressor blade leading edge toa mean radius point of the lowest pressure turbine rotor stage bladetrailing edge of the turbine. The core length may be in the range from150 cm to 350 cm, and optionally 160 cm to 320 cm. Optionally, forexample for an engine with a fan tip radius in the range from 95 cm to150 cm, the core length may be in the range from 160 cm to 260 cm, forexample 200 cm to 250 cm. Optionally, for example for an engine with afan tip radius in the range from 155 cm to 200 cm, the core length maybe in the range from 240 cm to 320 cm, for example 260 cm to 300 cm.

The ratio of the fan tip radius to the core length

$\left( \frac{{fan}\mspace{14mu}{tip}\mspace{14mu}{radius}}{{core}\mspace{14mu}{lengt}} \right)$may be in the range or 0.3 to 1, and optionally of 0.4 to 0.9,optionally 0.5 to 0.8. Optionally, for example for an engine with a fantip radius in the range from 95 cm to 150 cm, the ratio of the fan tipradius to the core length may be in the range from 0.4 to 0.9,optionally 0.5 to 0.8, or 0.55 to 0.75. Optionally, for example for anengine with a fan tip radius in the range from 155 cm to 200 cm, theratio of the fan tip radius to the core length may be in the range from0.5 to 0.8, optionally 0.60 to 0.80.

The gas turbine engine may further comprise a gearbox. The gearbox maybe connected between the core shaft and the fan. The gearbox may bearranged to receive an input from the core shaft, and to provide anoutput to drive the fan at a lower rotational speed than the core shaft.The gearbox may help to facilitate the ratios (for example the enginearea ratio) described and/or claimed herein.

The fan may comprise a plurality of fan blades extending radially from ahub, each fan blade having a leading edge and a trailing edge. Thelowest pressure turbine stage may comprise a row of rotor blades, eachof the rotor blades extending radially and having a leading edge and atrailing edge. The gas turbine engine may have a fan tip axis that joinsa radially outer tip of the leading edge of one of the plurality of fanblades and the radially outer tip of the trailing edge of one of therotor blades of the lowest pressure stage of the turbine. The fan tipaxis may lie in longitudinal plane which contains a centreline of thegas turbine engine. A fan tip axis angle may be defined as the anglebetween the fan tip axis and the centreline, and the fan axis angle isin a range between 10 to 20 degrees, optionally 11 to 18 degrees, 12 to17 degrees, or 12 to 16 degrees. The fan tip axis angle may be asdescribed below.

The engine core may comprise more than one turbine. The turbine may be afirst turbine, the compressor may be a first compressor, and the coreshaft may be a first core shaft. The engine core may further comprise asecond turbine, a second compressor, and a second core shaft connectingthe second turbine to the second compressor. The second turbine, secondcompressor, and second core shaft may be arranged to rotate at a higherrotational speed than the first core shaft.

According to another aspect, there is provided a gas turbine engine foran aircraft comprising an engine core comprising a turbine, acompressor, a core shaft connecting the turbine to the compressor, and acore exhaust nozzle having a core exhaust nozzle exit, the core exhaustnozzle having a core exhaust nozzle pressure ratio calculated usingtotal pressure at the core nozzle exit; a fan located upstream of theengine core, the fan comprising a plurality of fan blades; and a nacellesurrounding the fan and the engine core and defining a bypass ductlocated radially outside of the engine core, the bypass duct comprisinga bypass exhaust nozzle having a bypass exhaust nozzle exit, the bypassexhaust nozzle having a bypass exhaust nozzle pressure ratio calculatedusing total pressure at the bypass nozzle exit. A bypass to core ratioof:

$\frac{{bypass}\mspace{14mu}{exhaust}\mspace{14mu}{nozzle}\mspace{14mu}{pressure}\mspace{14mu}{ratio}}{{core}\mspace{14mu}{exhaust}\mspace{14mu}{nozzle}\mspace{14mu}{{pressur}e}\mspace{14mu}{ratio}}$is configured to be in the range from 1.1 to 2 under aircraft cruiseconditions.

The present aspect relates to a gas turbine engine with specifiedrelative core and bypass exhaust nozzle pressure ratios. The skilledperson will appreciate that a nozzle pressure ratio (NPR) is defined as:

$\frac{{total}\mspace{14mu}{pressure}\mspace{14mu}{at}\mspace{14mu}{nozzle}\mspace{14mu}{exit}}{{ambient}\mspace{14mu}{pressure}\mspace{14mu}{of}\mspace{14mu}{surroundings}}$

The skilled person will appreciate that, as is standard in the field,“total pressure” at a nozzle exit is defined as the sum of the staticand dynamic pressures at the nozzle exit. Given that the ambientpressure of the surroundings is equal for the core exhaust nozzle andthe bypass exhaust nozzle, the bypass to core ratio therefore may besimplified as follows:

$\frac{{NPR}_{{bypass}\mspace{14mu}{exhaust}\mspace{14mu}{nozzle}}}{{NPR}_{{core}\mspace{14mu}{exhau}\mspace{14mu}{nozzle}}} = {\frac{{total}\mspace{14mu}{pressure}\mspace{14mu}{at}\mspace{14mu}{bypass}\mspace{14mu}{nozzle}\mspace{14mu}{{exit}/{ambient}}\mspace{14mu}{pressure}}{{total}\mspace{14mu}{pressure}\mspace{14mu}{at}\mspace{14mu}{core}\mspace{14mu}{nozzle}\mspace{14mu}{{exit}/{ambient}}\mspace{14mu}{pressure}} = \frac{{total}\mspace{14mu}{pressure}\mspace{14mu}{at}\mspace{14mu}{bypass}\mspace{14mu}{nozzle}\mspace{14mu}{exit}}{{total}{\mspace{11mu}\;}{pressure}\mspace{14mu}{at}\mspace{14mu}{nozzle}\mspace{14mu}{exit}}}$

The ratio of

$\frac{{total}\mspace{14mu}{pressure}\mspace{14mu}{at}\mspace{14mu}{bypass}\mspace{14mu}{nozzle}\mspace{14mu}{exit}}{{total}\mspace{14mu}{pressure}\mspace{14mu}{at}\mspace{14mu}{core}\mspace{14mu}{nozzle}\mspace{14mu}{exit}}$may also be referred to as the extraction ratio. The ambient pressure(or pressure of the surroundings) may also be referred to as exit staticpressure.

The skilled person will appreciate that the specified relationshipbetween NPRs may improve engine efficiency as compared to known turbineengines, for example by improving fuel-burn. The skilled person willappreciate that the specified relationship between NPRs may allowdimension limitations for the engine core and/or the fan to beinferred—the specified relationship therefore is not limited in itsapplication to when the aircraft is at cruise conditions; rather,whether or not an engine falls within the scope of the claims may beinferred from those dimensions when the aircraft/engine is not in use.

The bypass to core ratio may be higher than that of known aircraft gasturbine engines.

The bypass to core ratio may be in the range from 1.1 to 2.0 underaircraft cruise conditions. The bypass to core ratio may be in the rangefrom 1.10 to 2.00 under aircraft cruise conditions. The bypass to coreratio may be above 1.15 under aircraft cruise conditions. The bypass tocore ratio may be in the range from 1.2 to 1.5 under aircraft cruiseconditions. The bypass to core ratio may be in the range from 1.1 to 1.6under aircraft cruise conditions. Optionally, for example for an enginewith a fan tip radius in the range from 110 cm to 150 cm, the bypass tocore ratio may be in the range from 1.0 to 1.4; for example from 1.1 to1.4 or from 1.0 to 1.3. Optionally, for example for an engine with a fantip radius in the range from 155 cm to 200 cm, the bypass to core ratiomay be in the range from 1.3 to 1.6.

Optionally, for example for an engine with a fan tip radius in the rangefrom 110 cm to 150 cm, NPR_(bypass exhaust nozzle) may be in the rangefrom 2.0 to 2.3. Optionally, for example for an engine with a fan tipradius in the range from 155 cm to 200 cm, NPR_(bypass exhau nozzle) maybe in the range from 2.1 to 2.3.

Optionally, for example for an engine with a fan tip radius in the rangefrom 110 cm to 150 cm, NPR_(core exhau nozzle) may be in the range from1.7 to 1.9. Optionally, for example for an engine with a fan tip radiusin the range from 155 cm to 200 cm, NPR_(core exhaust nozzle) may be inthe range from 1.4 to 1.6.

The bypass ratio is defined as the ratio of the mass flow rate of theflow through the bypass duct to the mass flow rate of the flow throughthe core at cruise conditions. The bypass ratio may be greater than (oron the order of) any of the following: 8, 8.5, 9, 9.5, 10, 10.5, 11,11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18,18.5, 19, 19.5, or 20. The bypass ratio may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The bypass ratio may be in therange from 11 to 20, and optionally in the range from 13 to 20 or 14 to20. The bypass ratio may be in the range of 8 to 9.5, for example forsome direct drive engines (engines without a gearbox). The bypass ratiomay be in the range of 9 to 16, for example for some geared engines(engines with a gearbox). Optionally, for example for an engine with afan tip radius in the range from 110 cm to 150 cm (which may be geared),the bypass ratio may be in the range from 9 to 15, and optionally from13 to 15. Optionally, for example for an engine with a fan tip radius inthe range from 155 cm to 200 cm (which may be geared), the bypass tocore ratio may be in the range from 13 to 18, optionally 13 to 16.

Cruise conditions may correspond to atmospheric conditions at analtitude that is in the range of from 10500 m to 11600 m, and optionallyto atmospheric conditions at an altitude of 11000 m. A forward speed ofthe gas turbine engine at cruise conditions may be in the range of fromMach number (Mn) 0.75 to Mn 0.85. A forward speed of the gas turbineengine at the cruise conditions may be Mn 0.8. Cruise conditions maycorrespond to:

-   -   a forward Mach number of 0.8;    -   a pressure of 23000 Pa; and    -   a temperature of −55 deg C.

The total pressure at the bypass nozzle exit may be determined at anexit plane of the bypass exhaust nozzle. The exit plane may extend froma rearmost point of the nacelle towards a centreline of the engine. Theexit plane may be a radial plane.

The engine core may comprise a casing (also referred to as an innerfixed structure). The total pressure at the core nozzle exit may bedetermined at an exit plane of the core exhaust nozzle. The exit planemay extend from a rearmost point of the engine core casing towards acentreline of the engine. The exit plane may be a radial plane. Theouter diameter of the bypass exhaust nozzle at the bypass exhaust nozzleexit may be in the range of 200 cm to 400 cm, and optionally 200 cm to380 cm. Optionally, for example for an engine with a fan tip radius inthe range from 95 cm to 150 cm (for example 110 cm to 150 cm), the outerdiameter of the bypass exhaust nozzle may be in the range from 200 cm to290 cm. Optionally, for example for an engine with a fan tip radius inthe range from 155 cm to 200 cm, the outer diameter of the bypassexhaust nozzle may be in the range from 290 cm to 380 cm.

The inner diameter of the bypass exhaust nozzle may be measured at theaxial position of the rearmost tip of the nacelle. The inner diameter ofthe bypass exhaust nozzle may be the radial distance between the outersurfaces of the engine core at the axial position of the rearmost tip ofthe nacelle. The inner diameter of the bypass exhaust nozzle may be inthe range from 100 cm to 250 cm, and optionally from 130 cm to 220 cm.Optionally, for example for an engine with a fan tip radius in the rangefrom 95 cm to 150 cm (for example 110 cm to 150 cm), the inner diameterof the bypass exhaust nozzle may be in the range from 130 cm to 180 cm.Optionally, for example for an engine with a fan tip radius in the rangefrom 155 cm to 200 cm, the inner diameter of the bypass exhaust nozzlemay be in the range from 160 cm to 220 cm.

The flow area of the core exhaust nozzle at the core exhaust nozzle exitmay be from 0.4 m² (600 square inches) to 1.3 m² (2000 square inches).Optionally, for example for an engine with a fan tip radius in the rangefrom 95 cm to 150 cm (for example 110 cm to 150 cm), the flow area ofthe core exhaust nozzle at the core exhaust nozzle exit may be in therange from 0.4 m² (600 square inches) to 0.6 m² (900 square inches).Optionally, for example for an engine with a fan tip radius in the rangefrom 155 cm to 200 cm, the flow area of the core exhaust nozzle at thecore exhaust nozzle exit may be in the range from 0.6 m² (900 squareinches) to 1.3 m² (2000 square inches).

The flow area of the bypass exhaust nozzle at the bypass duct exhaustnozzle exit may be from 1.9 m² (3000 square inches) to 5.8 m² (9000square inches). Optionally, for example for an engine with a fan tipradius in the range from 95 cm to 150 cm (for example 110 cm to 150 cm),the flow area of the bypass duct exhaust nozzle at the bypass ductnozzle exit may be in the range from 1.9 m² (3000 square inches) to 4.5m² (7000 square inches). Optionally, for example for an engine with afan tip radius in the range from 155 cm to 200 cm, the flow area of thebypass duct exhaust nozzle at the bypass duct exhaust nozzle exit may bein the range from 4.5 m² (7000 square inches) to 5.8 m² (9000 squareinches).

A ratio of bypass exhaust nozzle flow area to the core exhaust nozzleflow area

$\left( \frac{{bypass}\mspace{14mu}{exhaust}\mspace{14mu}{nozzle}\mspace{14mu}{flow}\mspace{14mu}{area}}{{core}\mspace{14mu}{exhaust}\mspace{14mu}{nozzle}\mspace{14mu}{flow}\mspace{14mu}{area}} \right)$may be in the range from 4 to 6, and optionally in the range from 5 to6. Optionally, a geared engine (with a gearbox) may have a ratio ofbypass exhaust nozzle flow area to the core exhaust nozzle flow area inthe range from 5 to 6.

The bypass exhaust nozzle and/or the core exhaust nozzle may be aconvergent nozzle.

The gas turbine engine may further comprise a gearbox connected betweenthe core and the fan. The gearbox may be arranged to receive an inputfrom the core shaft and to provide an output to drive the fan at a lowerrotational speed than the core shaft.

The fan tip radius, measured between a centreline of the engine and anoutermost tip of each fan blade at its leading edge, may be in the rangefrom 95 cm to 200 cm, for example in the range from 110 cm to 150 cm, oralternatively in the range from 155 cm to 200 cm. The fan tip radius maybe greater than any of: 110 cm, 115 cm, 120 cm, 125 cm, 130 cm, 135 cm,140 cm, 145 cm, 150 cm, 155 cm, 160 cm, 165 cm, 170 cm, 175 cm, 180 cm,185 cm, 190 cm or 195 cm. The fan tip radius may be around 110 cm, 115cm, 120 cm, 125 cm, 130 cm, 135 cm, 140 cm, 145 cm, 150 cm, 155 cm, 160cm, 165 cm, 170 cm, 175 cm, 180 cm, 185 cm, 190 cm or 195 cm. The fantip radius may be greater than 160 cm. The fan tip radius may be in therange from 95 cm to 150 cm, optionally in the range from 110 cm to 150cm, optionally in the range of from 110 cm to 145 cm, and furtheroptionally in the range from 120 cm to 140 cm. The fan tip radius may bein the range from 155 cm to 200 cm, optionally in the range from 160 cmto 200 cm, and further optionally in the range from 165 cm to 190 cm.

The engine core may comprise more than one turbine. The turbine may be afirst turbine, the compressor may be a first compressor, and the coreshaft may be a first core shaft. The engine core may comprise a secondturbine, a second compressor, and a second core shaft connecting thesecond turbine to the second compressor. The second turbine, secondcompressor, and second core shaft may be arranged to rotate at a higherrotational speed than the first core shaft.

The gas turbine engine may be arranged to be mounted beneath a wing ofan aircraft. A downstream blockage ratio may be defined as:

$\frac{\begin{matrix}{{the}\mspace{14mu}{turbine}\mspace{14mu}{diameter}\mspace{14mu}{at}\mspace{14mu}{an}\mspace{14mu}{axial}\mspace{14mu}{location}} \\{{{of}\mspace{14mu}{the}\mspace{14mu}{{turbine}'}s\mspace{14mu}{lowest}\mspace{11mu}{pressure}\mspace{14mu}{rotor}\mspace{14mu}{stage}}\;}\end{matrix}}{{ground}\mspace{14mu}{plane}\mspace{14mu}{to}\mspace{14mu}{wing}\mspace{14mu}{distance}}$A quasi-non-dimensional mass flow rate, Q, may be defined as:

$Q = {W\frac{\sqrt{T0}}{P{0.A_{flow}}}}$where:W is mass flow rate through the fan in Kg/s;T0 is average stagnation temperature of the air at the fan face inKelvin;P0 is average stagnation pressure of the air at the fan face in Pa; andA_(flow) is the flow area of the fan face in m².A Q ratio of:

-   -   the downstream blockage ratio×quasi non dimensional mass flow        rate Q may be in a range from 0.005 Kgs⁻¹N⁻¹K^(1/2) to 0.011        Kgs⁻¹N⁻¹K^(1/2).

The Q ratio may be in a range from 0.005 Kgs⁻¹N⁻¹K^(1/2) to 0.010Kgs⁻¹N⁻¹K^(1/2), and optionally from 0.0050 Kgs⁻¹N⁻¹K^(1/2) to 0.0110 orto 0.0100 Kgs⁻¹N⁻¹K^(1/2). The Q ratio may be in a range from 0.006Kgs⁻¹N⁻¹K^(1/2) to 0.009 Kgs⁻¹N⁻¹K^(1/2).

The downstream blockage ratio may be in a range from 0.2 to 0.3. Thedownstream blockage ratio may be in a range from 0.20 to 0.29. Thedownstream blockage ratio may be in a range from 0.22 to 0.28.

A specific thrust may be defined as net engine thrust divided by massflow rate through the engine; and at engine cruise conditions, it may bethat:0.029Kgs ⁻¹ N ⁻¹ K ^(1/2) ≤Q≤0.036Kgs ⁻¹ N ⁻¹ K ^(1/2); and70Nkg ⁻¹ s≤specific thrust≤110Nkg ⁻¹ s.

At cruise conditions, it may be that: 0.032 Kgs⁻¹N⁻¹K^(1/2)≤Q≤0.036Kgs⁻¹N⁻¹K^(1/2). At cruise conditions, it may be that: 0.033Kgs⁻¹N⁻¹K^(1/2)≤Q≤0.035 Kgs⁻¹N⁻¹K^(1/2).

According to another aspect, there is provided a method of operating anaircraft comprising a gas turbine engine comprising:

an engine core comprising a turbine, a compressor, a core shaftconnecting the turbine to the compressor, and a core exhaust nozzlehaving a core exhaust nozzle exit, the core exhaust nozzle having a coreexhaust nozzle pressure ratio calculated using total pressure at thecore nozzle exit;

a fan located upstream of the engine core, the fan comprising aplurality of fan blades; and

a nacelle surrounding the fan and the engine core and defining a bypassduct located radially outside of the engine core, the bypass ductcomprising a bypass exhaust nozzle having a bypass exhaust nozzle exit,the bypass exhaust nozzle having a bypass exhaust nozzle pressure ratiocalculated using total pressure at the bypass nozzle exit.

The method comprises controlling the aircraft such that a bypass to coreratio of:

$\frac{{bypass}\mspace{14mu}{exhaust}\mspace{14mu}{nozzle}\mspace{14mu}{pressure}\mspace{14mu}{ratio}}{{core}\mspace{14mu}{exhaust}\mspace{14mu}{nozzle}\mspace{14mu}{pressure}\mspace{14mu}{ratio}}$is in the range from 1.1 to 2 under aircraft cruise conditions.

The aircraft may be controlled such that a Q ratio of:

-   -   the downstream blockage ratio×quasi non dimensional mass flow        rate Q, as defined for the preceding aspect, may be in a range        from 0.005 Kgs⁻¹N⁻¹K^(1/2) to 0.011 Kgs⁻¹N⁻¹K^(1/2).

The Q ratio may be in a range from 0.005 Kgs⁻¹N⁻¹K^(1/2) to 0.010Kgs⁻¹N⁻¹K^(1/2), and optionally from 0.0050 Kgs⁻¹N⁻¹K^(1/2) to 0.0100Kgs⁻¹N⁻¹K^(1/2). The Q ratio may be in a range from 0.006Kgs⁻¹N⁻¹K^(1/2) to 0.009 Kgs⁻¹N⁻¹K^(1/2).

The gas turbine engine may be as described for the preceding aspect.

According to another aspect, there is provided a gas turbine engine foran aircraft comprising: an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor; afan located upstream of the engine core, the fan comprising a pluralityof fan blades extending from a hub; and a gearbox that receives an inputfrom the core shaft and outputs drive to the fan so as to drive the fanat a lower rotational speed than the core shaft, wherein: the gasturbine engine has an engine length and a centre of gravity positionmeasured relative to the fan, and wherein a centre of gravity positionratio of:

-   -   the centre of gravity position/the engine length        is in a range from 0.43 to 0.6.

Defining the centre of gravity position ratio in this range may allowthe centre of gravity to be located closer the front mounting positionof the gas turbine engine. This may help to reduce or minimise mountingloads compared to centre of gravity position ratios found in known gasturbine engines or which would be achieved with a proportional scalingof engine architecture. Other effects such as reducing bending of theengine core and deflection of the shaft may also be provided by definingthe centre of gravity position ratio as defined above. The centre ofgravity position ratio may be higher than that of known aircraft gasturbine engines.

The centre of gravity position ratio may be in a range from 0.43 to0.60, and optionally from 0.45 to 0.6 or 0.46 to 0.6 (or to 0.60). Thecentre of gravity position ratio may be in a range from 0.47 to 0.49,for example for an engine with a fan tip radius in the range from 110 cmto 150 cm. The centre of gravity position ratio may be in range from0.45 to 0.48, for example for an engine with a fan tip radius in therange from 155 cm to 200 cm.

The engine length may be in the range from 200 cm to 500 cm, andoptionally from 230 cm to 470 cm, optionally 300 cm to 450 cm.Optionally, for example for an engine with a fan tip radius in the rangefrom 110 cm to 150 cm (for example 120 cm to 140 cm), the engine lengthmay be in the range from 230 cm to 370 cm, optionally 300 to 360 cm.Optionally, for example for an engine with a fan tip radius in the rangefrom 155 cm to 200 cm (for example 165 cm to 190 cm), the engine lengthmay be in the range from 370 cm to 470 cm, optionally 390 cm to 450 cm.

The centre of gravity position may be in a range between 100 cm to 230cm, optionally 140 cm to 220 cm. Optionally, for example for an enginewith a fan tip radius in the range from 110 cm to 150 cm, the centre ofgravity position may be in a range between 100 cm to 180 cm, optionally140 cm to 180 cm. Optionally, for example for an engine with a fan tipradius in the range from 155 cm to 200 cm, the centre of gravityposition may be in a range between 160 cm to 230 cm, optionally 180 cmto 220 cm.

The engine length may be measured as the axial distance between aforward region of the fan and a rearward region of the turbine.

The turbine may comprise a lowest pressure turbine stage having a row ofrotor blades, and the engine length may be measured as the axialdistance between: the intersection of the leading edge of one of theplurality of fan blades and the hub; and a mean radius point of thetrailing edge of one of the rotor blades of the lowest pressure turbinestage of the turbine. The mean radius point may be the midpoint betweena 0% span position and a 100% span position of the rotor blade.

The turbine may be a lowest pressure turbine of a plurality of turbinesprovided in the core.

The position of centre of gravity may be measured as the axial distancebetween the intersection of the leading edge of one of the plurality offan blades and the hub; and the centre of gravity of the gas turbineengine.

According to an aspect, there is provided a fan speed to centre ofgravity ratio of:

-   -   the centre of gravity position ratio×maximum take off rotational        fan speed        may be in a range from 600 rpm to 1350 rpm. This ratio may be        lower than that of known aircraft gas turbine engines. By        defining the fan speed to centre of gravity ratio in this range        the centre of gravity may be moved forwards relative to that of        a direct drive engine whilst also providing a relatively low fan        rotational speed.

The fan speed to centre of gravity ratio may be in a range from 650 rpmto 1276 rpm. The fan speed to centre of gravity ratio may be in a rangefrom 600 rpm to 1290 rpm. Optionally, for example for an engine with afan tip radius in the range from 110 cm to 150 cm, the fan speed tocentre of gravity ratio may be 925 rpm to 1325 rpm. Optionally, forexample for an engine with a fan tip radius in the range from 155 cm to200 cm, the fan speed to centre of gravity ratio may be 650 rpm to 910rpm.

The maximum take-off rotational fan speed may be in a range between 1450rpm to 3020 rpm. Optionally, for example for an engine with a fan tipradius in the range from 110 cm to 150 cm, the maximum take-offrotational fan speed may be in a range between 1970 rpm to 3020 rpm.Optionally, for example for an engine with a fan tip radius in the rangefrom 155 cm to 200 cm, the maximum take-off rotational fan speed may bein a range between 1450 rpm to 1910 rpm.

According to another aspect, there is provided a method of operating anaircraft comprising a gas turbine engine comprising: an engine corecomprising a turbine, a compressor, and a core shaft connecting theturbine to the compressor; a fan located upstream of the engine core,the fan comprising a plurality of fan blades extending from a hub; and agearbox that receives an input from the core shaft and outputs drive tothe fan so as to drive the fan at a lower rotational speed than the coreshaft, wherein: the gas turbine engine has an engine length and a centreof gravity position measured relative to the fan, and wherein the methodcomprises controlling the aircraft such that a centre of gravityposition ratio of:

-   -   the centre of gravity position/the engine length        is in a range from 0.43 to 0.6, and a fan speed to centre of        gravity ratio of:    -   the centre of gravity position ratio×maximum take off rotational        fan speed        is in a range from 600 rpm to 1350 rpm.

The fan speed to centre of gravity ratio may be in a range from 650 rpmto 1276 rpm. Optionally, for example for an engine with a fan tip radiusin the range from 110 cm to 150 cm, the fan speed to centre of gravityratio may be 925 rpm to 1325 rpm. Optionally, for example for an enginewith a fan tip radius in the range from 155 cm to 200 cm, the fan speedto centre of gravity ratio may be 650 rpm to 910 rpm.

The maximum take-off rotational fan speed may be in a range between 1450rpm to 3020 rpm. Optionally, for example for an engine with a fan tipradius in the range from 110 cm to 150 cm, the maximum take-offrotational fan speed may be in a range between 1970 rpm to 3020 rpm.Optionally, for example for an engine with a fan tip radius in the rangefrom 155 cm to 200 cm, the maximum take-off rotational fan speed may bein a range between 1450 rpm to 1910 rpm.

According to another aspect, there is provided a gas turbine engine foran aircraft comprising: an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor; afan located upstream of the engine core, the fan comprising a pluralityof fan blades extending from a hub; and a gearbox that receives an inputfrom the core shaft and outputs drive to the fan so as to drive the fanat a lower rotational speed than the core shaft, wherein: the gasturbine engine has an engine length and a gearbox location relative to aforward region of the fan, wherein a gearbox location ratio of:

-   -   gearbox location/engine length        is in a range from 0.19 to 0.45.

Defining gearbox location ratio in this range may allow the gearbox tobe located at or near a front mounting position of the gas turbineengine. As the gearbox is generally amongst the heaviest componentswithin the engine its location may have a significant influence on theposition of the centre of gravity. Moving the centre of gravity closerto the front mounting may help to minimise rear mounting loads. Othereffects such as reducing bending of the engine core and deflection ofthe core connecting drive shaft may also be provided by controlling theengine centre of gravity by suitable positioning of the gearbox. Thegearbox location ratio may be higher than that of known gas turbineengines.

The gearbox location ratio may be in a range from 0.19 to 0.3. Thegearbox location ratio may be in a range from 0.19 to 0.23. The gearboxlocation ratio may be in a range from 0.19 to 0.23, for example 0.19 to0.21, for example for an engine with a fan tip radius in the range from110 cm to 150 cm. The gearbox location ratio may be in a range from 0.20to 0.25, for example equal to or around 0.23—for example, being in therange from 0.225 to 0.235—for example, for an engine with a fan tipradius in the range from 155 cm to 200 cm.

The engine length may be in the range from 200 cm to 500 cm, andoptionally from 230 cm to 470 cm, optionally 300 cm to 450 cm.Optionally, for example for an engine with a fan tip radius in the rangefrom 110 cm to 150 cm (for example 120 cm to 140 cm), the engine lengthmay be in the range from 230 cm to 370 cm, optionally 300 to 360 cm.Optionally, for example for an engine with a fan tip radius in the rangefrom 155 cm to 200 cm (for example 165 cm to 190 cm), the engine lengthmay be in the range from 370 cm to 470 cm, optionally 390 cm to 450 cm.

The gearbox location may be in a range between 50 cm to 110 cm.Optionally, for example for an engine with a fan tip radius in the rangefrom 110 cm to 150 cm, the gearbox location may be in a range between 50cm to 80 cm, optionally 55 cm to 75 cm. Optionally, for an engine with afan tip radius in the range from 155 cm to 200 cm, the gearbox locationmay be in a range between 80 cm to 110 cm, optionally 85 cm to 105 cm.

The engine length may be measured as the axial distance between aforward region of the fan and a rearward region of the turbine.

The turbine may comprise a lowest pressure turbine stage having a row ofrotor blades, and the engine length is measured as the axial distancebetween: the intersection of the leading edge of one of the plurality offan blades and the hub; and a mean radius point of the trailing edge ofone of the rotor blades of the lowest pressure turbine stage of theturbine. The mean radius point may be the midpoint between a 0% spanposition and a 100% span position of the rotor blade.

The turbine may be a lowest pressure turbine of a plurality of turbinesprovided in the core.

The gearbox location may be measured between: the intersection of aleading edge of one of the fan blades and the hub; and a radial centreplane of the gearbox, the radial centre plane being at the midpointbetween the front face of a most forward gear mesh of the gearbox andthe rear face of a most rearward gear mesh of the gearbox.

The gearbox may be an epicyclic gearbox comprising a ring gear; in suchembodiments, the gearbox location may be measured as the axial distancebetween: the intersection of a leading edge of one of the fan blades andthe hub; and a radial plane intersecting the axial centre point of thering gear.

The fan blades may be formed at least partly from a composite material,and the gearbox location for such embodiments may be in a range between50 cm and 110 cm and optionally in a range between 80 cm and 110 cm.

The fan blades may be formed at least partly from a metal or metalalloy, such as an aluminium-lithium alloy, and wherein the gearboxlocation may be in a range between 50 cm and 110 cm and optionally in arange between 50 cm and 80 cm.

The fan blades may be formed at least partly from a composite materialand the gearbox location ratio may be in the range of from 0.02 to 0.25,for example equal to or around 0.23; for example, being in the rangefrom 0.225 to 0.235. This may be, for example, for an engine with a fantip radius in the range from 155 cm to 200 cm.

The fan blades may be formed at least partly from a metal or metalalloy, such as an aluminium-lithium alloy, and the gearbox locationratio may be in a range from 0.19 to 0.25, for example 0.19 to 0.23.This may be, for example, for an engine with any fan tip radius.

According to another aspect, there is provided a gas turbine engine foran aircraft comprising: an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor; afan located upstream of the engine core, the fan comprising a pluralityof fan blades extending from a hub, the fan blades being formed from ametal or metal alloy (optionally an aluminium-lithium alloy); and agearbox that receives an input from the core shaft and outputs drive tothe fan so as to drive the fan at a lower rotational speed than the coreshaft, wherein: the gas turbine engine has an engine length and agearbox location relative to a forward region of the fan, wherein agearbox location ratio of:

-   -   gearbox location/engine length        wherein the gearbox location ratio may be in a range from 0.19        to 0.3, optionally 0.19 to 0.25.

This may be, for example, for an engine with a fan tip radius in therange from 110 cm to 150 cm.

According to another aspect, there is provided a gas turbine engine foran aircraft comprising: an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor; afan located upstream of the engine core, the fan comprising a pluralityof fan blades extending from a hub, the fan blades being formed from atleast partly from a composite material; and a gearbox that receives aninput from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft, wherein: thegas turbine engine has an engine length and a gearbox location relativeto a forward region of the fan, wherein a gearbox location ratio of:

-   -   gearbox location/engine length        wherein the gearbox location ratio may be in the range of from        0.20 to 0.25, for example equal to or around 0.23; for example,        being in the range from 0.225 to 0.235.

This may be, for example, for an engine with a fan tip radius in therange from 155 cm to 200 cm.

The gas turbine engine of these aspects may further have a centre ofgravity position ratio as defined in any of the statements of theprevious two aspects.

According to another aspect there is provided a gas turbine engine foran aircraft comprising an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor; afan located upstream of the engine core, the fan comprising a pluralityof fan blades, wherein a fan tip radius of the fan is measured/definedbetween a centreline of the engine and an outermost tip of each fanblade at its leading edge; and a nacelle surrounding the fan and theengine core and defining a bypass exhaust nozzle, the bypass exhaustnozzle having an outer radius. An outer bypass to fan ratio of:

$\frac{{the}\mspace{14mu}{outer}\mspace{14mu}{radius}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{bypass}\mspace{14mu}{exhaust}\mspace{14mu}{nozzle}}{{the}\mspace{14mu}{fan}\mspace{14mu}{tip}\mspace{14mu}{radius}}$

is in the range from 0.6 to 1.05.

The present aspect relates to a gas turbine engine comprising a fan andnacelle with specified relative shapes and/or sizes. The skilled personwould appreciate that having a relatively narrow bypass exhaust nozzle,as compared to fan size, may reduce drag produced by the engine in use.Further, the skilled person would appreciate that the relatively narrowbypass exhaust nozzle may create a more compact exhaust system, whichmay allow or facilitate under-wing installation of a larger engine on anaircraft. The outer bypass to fan ratio may be lower than that of knownaircraft gas turbine engines.

The outer bypass to fan ratio may be in the range from 0.60 to 1.05. Theouter bypass to fan ratio may be in the range from 0.65 to 1.00. Theouter bypass to fan ratio may be lower than 1.05, optionally lower than1.02, and further optionally lower than 1.00. The outer bypass to fanratio may be in the range from 0.80 to 1.00. The outer bypass to fanratio may be in the range from 0.9 to 1.0, and optionally in the rangefrom 0.90 to 1.00.

The fan tip radius may be in the range from 95 cm to 200 cm, for examplein the range from 110 cm to 150 cm, or alternatively in the range from155 cm to 200 cm. Optionally, for example for an engine with a fan tipradius in the range from 110 cm to 150 cm, the outer bypass to fan ratiomay be in the range from 0.95 to 1.00, for example equal to or around0.97, for example being in the range from 0.96 to 0.98. Optionally, forexample for an engine with a fan tip radius in the range from 155 cm to200 cm, outer bypass to fan ratio may be in the range from 0.91 to 0.98,for example equal to or around 0.95, for example being in the range from0.94 to 0.96.

The bypass exhaust nozzle may have an exit plane, which may be a radialplane. The outer radius of the bypass exhaust nozzle may be measured atthe axial position of the exit plane of the bypass exhaust nozzle.

The outer radius of the bypass exhaust nozzle may be measured at theaxial position of the rearmost tip of the nacelle. The outer radius ofthe bypass exhaust nozzle may be the radial distance between thecentreline of the engine and an inner surface of the nacelle at theaxial position of the rearmost tip of the nacelle.

The outer radius of the bypass exhaust nozzle may be in the range of 100cm to 200 cm, and optionally 100 cm to 190 cm. The outer radius of thebypass exhaust nozzle may be defined as half the diameter of the bypassexhaust nozzle as described above. Optionally, for example for an enginewith a fan tip radius in the range from 110 cm to 150 cm, the outerradius of the bypass exhaust nozzle may be in the range from 100 cm to145 cm, for example 110 cm to 140 cm. Optionally, for example for anengine with a fan tip radius in the range from 155 cm to 200 cm, theouter radius of the bypass exhaust nozzle may be in the range from 145cm to 190 cm.

The bypass exhaust nozzle may have an inner radius. An inner bypass tofan ratio of:

$\frac{{the}\mspace{14mu}{inner}\mspace{14mu}{radius}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{bypass}\mspace{14mu}{exhaust}\mspace{14mu}{nozzle}}{{the}\mspace{14mu}{fan}\mspace{14mu}{tip}\mspace{14mu}{radius}}$

may be in the range from 0.4 to 0.65.

The inner bypass to fan ratio may be lower than that for known aircraftgas turbine engines. The inner bypass to fan ratio may be in the rangefrom 0.5 to 0.6, and optionally in the range from 0.50 to 0.60. Theinner bypass to fan ratio may be in the range from 0.40 to 0.65. Theinner bypass to fan ratio may be lower than 0.65, and optionally lowerthan 0.64, and optionally lower than 0.62. The inner bypass to fan ratiomay be in the range from 0.54 to 0.64. Optionally, for example for anengine with a fan tip radius in the range from 110 cm to 150 cm, theinner bypass to fan ratio may be in the range of from 0.57 to 0.63, forexample 0.57 to 0.62, for example around 0.59, for example being in therange from 0.58 to 0.60. Optionally, for example for an engine with afan tip radius in the range from 155 cm to 200 cm, the inner bypass tofan ratio may be in the range from 0.5 to 0.6, and optionally from 0.52to 0.58.

The bypass exhaust nozzle may have an exit plane, which may be a radialplane. The inner radius of the bypass exhaust nozzle may be measured atthe axial position of the exit plane of the bypass exhaust nozzle. Theinner and outer radii may therefore be measured in the same radialplane. The inner and outer radii of the bypass exhaust nozzle maytherefore be measured at the exit of the bypass exhaust nozzle.

The inner radius of the bypass exhaust nozzle may be measured at theaxial position of the rearmost tip of the nacelle. The inner radius ofthe bypass exhaust nozzle may be the radial distance between thecentreline of the engine and an outer surface of the engine core at theaxial position of the rearmost tip of the nacelle. The inner radius ofthe bypass exhaust nozzle is half the inner diameter of the bypassexhaust nozzle. The inner radius of the bypass exhaust nozzle may be inthe range from 50 cm to 125 cm, and optionally from 65 cm to 110 cm,optionally from 75 cm to 110 cm. Optionally, for example for an enginewith a fan tip radius in the range from 110 cm to 150 cm, the innerradius of the bypass exhaust nozzle may be in the range from 70 cm to 90cm. Optionally, for example for an engine with a fan tip radius in therange from 155 cm to 200 cm, the inner radius of the bypass exhaustnozzle may be in the range from 80 cm to 120 cm, for example 90 cm to120 cm.

The nacelle may define a bypass duct located radially outside of theengine core. The bypass exhaust nozzle may be an exit of the bypassduct, for example forming a rearmost portion of the bypass exhaust duct.The nacelle may comprise an inner surface at least partly defining thebypass duct.

The engine may further comprise a bypass duct outlet guide vaneextending radially across the bypass duct between an outer surface ofthe engine core and the inner surface of the nacelle. The bypass ductoutlet guide vane may extend between a radially inner tip and a radiallyouter tip and may have a leading edge and a trailing edge relative tothe direction of gas flow through the bypass duct. An outer wall axismay be defined joining the radially outer tip of the trailing edge ofthe bypass duct outlet guide vane and the rearmost tip of the innersurface of the nacelle. The outer wall axis may lie in a longitudinalplane containing the centreline of the gas turbine engine. An outerbypass duct wall angle may be defined as the angle between the outerwall axis and the centreline.

The outer bypass duct wall angle may be in a range between −15 degreesand 1 degrees. The outer bypass duct wall angle may be in a rangebetween −10 degrees and 0 degrees, for example −5 degrees and 0 degrees.The outer bypass duct wall angle may be in a range between −4 degreesand −1 degrees. The outer bypass duct wall angle may be in the range of−0.5 degrees to −4 degrees; this may be for an engine with a fan tipradius in the range from 110 cm to 150 cm. The outer bypass duct wallangle may be in a range between −1 degrees and −5 degrees, for example−2.5 degrees to −4 degrees; this may be for an engine with a fan tipradius in the range from 155 cm to 200 cm.

A negative value of the outer bypass duct wall angle may correspond tothe outer wall axis sloping towards the centreline of the gas turbineengine.

The gas turbine engine may comprise a gearbox that receives an inputfrom the core shaft and outputs drive to the fan so as to drive the fanat a lower rotational speed than the core shaft. The gearbox may have agear ratio in the ranges defined elsewhere herein, for example 3.2 to 5or 3.2 to 3.8.

The gas turbine engine may comprise more than one turbine. The turbinemay be a first turbine, the compressor may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

According to another aspect there is provided a gas turbine engine foran aircraft comprising an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor; afan located upstream of the engine core, the fan comprising a pluralityof fan blades, wherein a fan tip radius of the fan is measured between acentreline of the engine and an outermost tip of each fan blade at itsleading edge; and a nacelle surrounding the fan and the engine core anddefining a bypass exhaust nozzle, the bypass exhaust nozzle having aninner radius. An inner bypass to fan ratio of:

$\frac{{the}\mspace{14mu}{inner}\mspace{14mu}{radius}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{bypass}\mspace{14mu}{exhaust}\mspace{14mu}{nozzle}}{{the}\mspace{14mu}{fan}\mspace{14mu}{tip}\mspace{14mu}{radius}}$

is in the range from 0.4 to 0.65.

The present aspect relates to a gas turbine engine comprising a fan andengine core with specified relative shapes and/or sizes. The skilledperson would appreciate that having a relatively narrow bypass exhaustnozzle, as compared to fan size, may reduce drag produced by the enginein use. The skilled person would appreciate that the engine core issituated radially within the bypass exhaust nozzle, and that the innerradius of the bypass exhaust nozzle may therefore equivalently bethought of as an outer radius of the engine core. The skilled personwould appreciate that having a relatively narrow engine core, ascompared to fan size, may reduce drag produced by the engine in use. Theinner bypass to fan ratio may be lower than that for known aircraft gasturbine engines.

The inner bypass to fan ratio may be in the range from 0.40 to 0.65. Theinner bypass to fan ratio may be lower than 0.65 and optionally lowerthan 0.64. The inner bypass to fan ratio may be in the range from 0.54to 0.64. Optionally, for example for an engine with a fan tip radius inthe range from 110 cm to 150 cm, the inner bypass to fan ratio may be inthe range of from 0.57 to 0.63, for example 0.57 to 0.62, for examplearound 0.59, for example being in the range from 0.58 to 0.60.Optionally, for example for an engine with a fan tip radius in the rangefrom 155 cm to 200 cm, the inner bypass to fan ratio may be in the rangefrom 0.5 to 0.6, and optionally from 0.52 to 0.58, optionally 0.53 to0.55.

The bypass exhaust nozzle may have an exit plane, which may be a radialplane. The inner radius of the bypass exhaust nozzle may be measured atthe axial position of the exit plane of the bypass exhaust nozzle.

The inner radius of the bypass exhaust nozzle may be measured at theaxial position of the rearmost tip of the nacelle. The inner radius ofthe bypass exhaust nozzle may be the radial distance between thecentreline of the engine and an outer surface of the engine core at theaxial position of the rearmost tip of the nacelle.

The inner radius of the bypass exhaust nozzle may be in the range from50 cm to 125 cm, and optionally from 65 cm to 110 cm, optionally from 75cm to 110 cm. Optionally, for example for an engine with a fan tipradius in the range from 110 cm to 150 cm, the inner radius of thebypass exhaust nozzle may be in the range from 70 cm to 90 cm.Optionally, for example for an engine with a fan tip radius in the rangefrom 155 cm to 200 cm, the inner radius of the bypass exhaust nozzle maybe in the range from 80 cm to 120 cm, for example 90 cm to 120 cm.

The bypass exhaust nozzle may have an outer radius, and an outer bypassto fan ratio of:

$\frac{{the}\mspace{14mu}{outer}\mspace{14mu}{radius}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{bypass}\mspace{14mu}{exhaust}\mspace{14mu}{nozzle}}{{the}\mspace{14mu}{fan}\mspace{14mu}{tip}\mspace{14mu}{radius}}$

may be in the range from 0.6 to 1.05. The outer bypass to fan ratio maybe lower than that for known aircraft gas turbine engines.

The outer bypass to fan ratio may be in the range from 0.9 to 1.0, andoptionally from 0.90 to 1.00. The outer bypass to fan ratio may be lowerthan 1.05, optionally lower than 1.02, and further optionally lower than1.00. The outer bypass to fan ratio may be in the range from 0.8 to1.00, and optionally in the range from 0.90 to 1.00. Optionally, forexample for an engine with a fan tip radius in the range from 110 cm to150 cm, the outer bypass to fan ratio may be in the range from 0.95 to1.00, for example equal to or around 0.97, for example being in therange from 0.96 to 0.98. Optionally, for example for an engine with afan tip radius in the range from 155 cm to 200 cm, outer bypass to fanratio may be in the range from 0.91 to 0.98, for example equal to oraround 0.95, for example being in the range from 0.94 to 0.96.

The bypass exhaust nozzle may have an exit plane, which may be radial.The outer radius of the bypass exhaust nozzle may be measured at theaxial position of the exit plane of the bypass exhaust nozzle. The outerradius of the bypass exhaust nozzle may be measured at the axialposition of the rearmost tip of the nacelle. The outer radius of thebypass exhaust nozzle may be the radial distance between the centrelineof the engine and an inner surface of the nacelle at the axial positionof the rearmost tip of the nacelle.

The outer radius of the bypass exhaust nozzle may be in the range of 100cm to 200 cm, and optionally 100 cm to 190 cm. The outer radius of thebypass exhaust nozzle may be defined as half the diameter of the bypassexhaust nozzle as described above. Optionally, for example for an enginewith a fan tip radius in the range from 95 cm to 150 cm, the outerradius of the bypass exhaust nozzle may be in the range from 100 cm to145 cm. Optionally, for example for an engine with a fan tip radius inthe range from 155 cm to 200 cm, the outer radius of the bypass exhaustnozzle may be in the range from 145 cm to 190 cm.

The nacelle may define a bypass duct located radially outside of theengine core. The nacelle may comprise an inner surface at least partlydefining the bypass duct.

The bypass exhaust nozzle may be an exit of the bypass duct, for exampleforming a rearmost portion of the bypass exhaust duct.

The engine may further comprise a bypass duct outlet guide vaneextending radially across the bypass duct between an outer surface ofthe engine core and the inner surface of the nacelle. The bypass ductoutlet guide vane may extend between a radially inner tip and a radiallyouter tip and may have a leading edge and a trailing edge relative tothe direction of gas flow through the bypass duct. An outer wall axismay be defined joining the radially outer tip of the trailing edge ofthe bypass duct outlet guide vane and the rearmost tip of the innersurface of the nacelle. The outer wall axis may lie in a longitudinalplane containing the centreline of the gas turbine engine. An outerbypass duct wall angle may be defined as the angle between the outerwall axis and the centreline.

The outer bypass duct wall angle may be in a range between −15 to 1degrees. The outer bypass duct wall angle may be in a range between −10degrees and 0 degrees, for example −5 degrees and 0 degrees. The outerbypass duct wall angle may be in a range between −4 degrees and −1degrees. The outer bypass duct wall angle may be in the range of −0.5degrees to −4 degrees; this may be for an engine with a fan tip radiusin the range from 110 cm to 150 cm. The outer bypass duct wall angle maybe in a range between −1 degrees and −5 degrees, for example −2.5degrees to −4.0 degrees; this may be for an engine with a fan tip radiusin the range from 155 cm to 200 cm.

A negative value of the outer bypass duct wall angle may correspond tothe outer wall axis sloping towards the centreline of the gas turbineengine.

The gas turbine engine may further comprise a gearbox that receives aninput from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft. The gearbox mayhave a gear ratio in the ranges defined elsewhere herein, for example3.2 to 5 or 3.2 to 3.8.

The gas turbine engine may comprise more than one turbine. The turbinemay be a first turbine, the compressor may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft are arranged to rotateat a higher rotational speed than the first core shaft.

According to another aspect, there is provided a gas turbine engine foran aircraft comprising: an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor; afan located upstream of the engine core, the fan comprising a pluralityof fan blades; a nacelle surrounding the gas turbine engine, the nacellecomprising an inner surface at least partly defining a bypass ductlocated radially outside of the engine core; and a bypass duct outletguide vane (OGV) extending radially across the bypass duct between anouter surface of the engine core and the inner surface of the nacelle,wherein the bypass duct outlet guide vane extends between a radiallyinner tip and a radially outer tip and has a leading edge and a trailingedge relative to the direction of gas flow through the bypass duct, anouter wall axis is defined joining the radially outer tip of thetrailing edge of the bypass duct outlet guide vane and the rearmost tipof the inner surface of the nacelle, wherein the outer wall axis lies ina longitudinal plane containing the centreline of the gas turbineengine, an outer bypass duct wall angle is defined as the angle betweenthe outer wall axis and the centreline, and the outer bypass duct wallangle is in a range between −15 to 1 degrees.

By providing the bypass duct wall angle in this range a more compactexhaust system may be provided. By using the angle range above thelength of the nacelle between the bypass OGV and the rearmost inner tipof the nacelle inner wall may be reduced. This may provide a shorterlength of nacelle which may provide a reduction in drag compared toknown gas turbine engines, or which would be achieved if the nacelledimensions were scaled proportionately when increasing the size of thegas turbine engine. The bypass duct wall angle may be lower (i.e. morenegative) than that of known gas turbine engines.

The outer bypass duct wall angle may be in a range between −10 degreesand 0 degrees, for example −5 degrees and 0 degrees. The outer bypassduct wall angle may be in a range between −4 degrees and −1 degrees. Theouter bypass duct wall angle may be in the range of −0.5 degrees to −4degrees; this may be for an engine with a fan tip radius in the rangefrom 110 cm to 150 cm. The outer bypass duct wall angle may be in arange between −1 degrees and −5 degrees, for example −2.5 degrees to −4degrees; this may be for an engine with a fan tip radius in the rangefrom 155 cm to 200 cm.

A negative value of the outer bypass duct wall angle may correspond tothe outer wall axis sloping towards the centreline of the gas turbineengine.

A bypass duct outlet guide vane radius, measured radially between theengine centreline and the radially outer tip of the trailing edge of thebypass OGV, may be in a range from 90 cm to 210 cm. Optionally, forexample for an engine with a fan tip radius in the range from 110 cm to150 cm, the bypass duct outlet guide vane radius may be in the rangefrom 90 cm to 150 cm, optionally 110 cm to 135 cm. Optionally, forexample for an engine with a fan tip radius in the range from 155 cm to200 cm, the bypass duct outlet guide vane radius may be in the rangefrom 160 cm to 210 cm, optionally 170 cm to 200 cm.

The gas turbine engine may further comprise a gearbox that receives aninput from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft. The gearbox mayhave a gear ratio in the ranges defined elsewhere herein, for example3.2 to 5 or 3.2 to 3.8.

The rearmost inner tip of the nacelle inner wall may be movable in useof the gas turbine engine to provide a variable area bypass duct exhaustnozzle (also referred to as a fan nozzle). The outer wall axis may bedefined based on the position of the rearmost tip of the inner surfaceof the nacelle during cruise conditions. Cruise conditions may be asdefined in connection with any other aspect.

The gas turbine engine of this aspect may further have an outer bypassto fan ratio and/or inner bypass to fan ratio as defined in the previousrelevant aspects of any of the above statements.

A fan tip radius of the fan may be measured between a centreline of theengine and an outermost tip of each fan blade at its leading edge; andthe nacelle may surround the fan and the engine core and define a bypassexhaust nozzle, the bypass exhaust nozzle having an outer radius.

An outer bypass to fan ratio of:

$\frac{{the}\mspace{14mu}{outer}\mspace{14mu}{radius}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{bypass}\mspace{14mu}{exhaust}\mspace{14mu}{nozzle}}{{the}\mspace{14mu}{fan}\mspace{14mu}{tip}\mspace{14mu}{radius}}$

may be in the range from 0.6 to 1.05. The outer bypass to fan ratio maybe in the range from 0.60 to 1.05. The outer bypass to fan ratio may bein the range from 0.65 to 1.00. The outer bypass to fan ratio may belower than 1.05, optionally lower than 1.02, and further optionallylower than 1.00. The outer bypass to fan ratio may be in the range from0.80 to 1.00. The outer bypass to fan ratio may be in the range from 0.9to 1.0, and optionally in the range from 0.90 to 1.00. Optionally, forexample for an engine with a fan tip radius in the range from 110 cm to150 cm, the outer bypass to fan ratio may be in the range 0.95 to 1.00,optionally equal to or around 0.97, for example being in the range from0.96 to 0.98. Optionally, for example for an engine with a fan tipradius in the range from 155 cm to 200 cm, outer bypass to fan ratio maybe in the range 0.91 to 0.98, optionally equal to or around 0.95, forexample being in the range from 0.94 to 0.96. The outer bypass to fanratio may be lower than that for known aircraft gas turbine engines.

The bypass exhaust nozzle may have an exit plane, which may be a radialplane. The outer radius of the bypass exhaust nozzle may be measured atthe axial position of the exit plane of the bypass exhaust nozzle. Theouter radius of the bypass exhaust nozzle may be measured at the axialposition of the rearmost tip of the nacelle.

The outer radius of the bypass exhaust nozzle may be the radial distancebetween the centreline of the engine and an inner surface of the nacelleat the axial position of the rearmost tip of the nacelle. The outerradius of the bypass exhaust nozzle may be in the range of 100 cm to 200cm, and optionally 100 cm to 190 cm. The outer radius of the bypassexhaust nozzle may be defined as half the diameter of the bypass exhaustnozzle as described above. Optionally, for example for an engine with afan tip radius in the range from 95 cm to 150 cm, the outer radius ofthe bypass exhaust nozzle may be in the range from 100 cm to 145 cm.Optionally, for example for an engine with a fan tip radius in the rangefrom 155 cm to 200 cm, the outer radius of the bypass exhaust nozzle maybe in the range from 145 cm to 190 cm.

The bypass exhaust nozzle may have an inner radius. An inner bypass tofan ratio of:

$\frac{{the}\mspace{14mu}{inner}\mspace{14mu}{radius}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{bypass}\mspace{14mu}{exhaust}\mspace{14mu}{nozzle}}{{the}\mspace{14mu}{fan}\mspace{14mu}{tip}\mspace{14mu}{radius}}$may be in the range from 0.4 to 0.65. The inner bypass to fan ratio maybe lower than that for known aircraft gas turbine engines. The innerbypass to fan ratio may be in the range from 0.5 to 0.6, and optionallyin the range from 0.50 to 0.60. The inner bypass to fan ratio may be inthe range from 0.40 to 0.65. The inner bypass to fan ratio may be lowerthan 0.65, and optionally lower than 0.64, and optionally lower than0.62. The inner bypass to fan ratio may be in the range from 0.54 to0.64. Optionally, for example for an engine with a fan tip radius in therange from 110 cm to 150 cm, the inner bypass to fan ratio may be in therange of from 0.57 to 0.63, for example 0.57 to 0.62, for example around0.59, for example being in the range from 0.58 to 0.60. Optionally, forexample for an engine with a fan tip radius in the range from 155 cm to200 cm, the inner bypass to fan ratio may be in the range from 0.5 to0.6, and optionally from 0.52 to 0.58, optionally 0.53 to 0.55.

The bypass exhaust nozzle may have an exit plane, which may be a radialplane. The inner radius of the bypass exhaust nozzle may be measured atthe axial position of the exit plane of the bypass exhaust nozzle. Theinner and outer radii may therefore be measured in the same radialplane. The inner and outer radii of the bypass exhaust nozzle maytherefore be measured at the exit of the bypass exhaust nozzle.

The inner radius of the bypass exhaust nozzle may be measured at theaxial position of the rearmost tip of the nacelle. The inner radius ofthe bypass exhaust nozzle may be the radial distance between thecentreline of the engine and an outer surface of the engine core at theaxial position of the rearmost tip of the nacelle. The inner radius ofthe bypass exhaust nozzle may be in the range from 50 cm to 125 cm, andoptionally from 65 cm to 110 cm, optionally from 75 cm to 110 cm.Optionally, for example for an engine with a fan tip radius in the rangefrom 110 cm to 150 cm, the inner radius of the bypass exhaust nozzle maybe in the range from 70 cm to 90 cm. Optionally, for example for anengine with a fan tip radius in the range from 155 cm to 200 cm, theinner radius of the bypass exhaust nozzle may be in the range from 80 cmto 120 cm, for example 90 cm to 120 cm.

According to another aspect, there is provided a gas turbine engine foran aircraft comprising: an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor; afan located upstream of the engine core, the fan comprising a pluralityof fan blades extending radially from a hub, each fan blade having aleading edge and a trailing edge; and wherein: the turbine comprises alowest pressure turbine stage having a row of rotor blades, each of therotor blades extending radially and having a leading edge and a trailingedge; the gas turbine engine has a fan tip axis that joins: a radiallyouter tip of the leading edge of one of the plurality of fan blades; andthe radially outer tip of the trailing edge of one of the rotor bladesof the lowest pressure stage of the turbine, the fan tip axis lying in alongitudinal plane which contains a centreline of the gas turbineengine, and a fan tip axis angle is defined as the angle between the fantip axis and the centreline, and the fan tip axis angle is in a rangebetween 10 to 20 degrees.

By providing a fan tip axis angle (also referred to as a fan axis angle)in this range the gas turbine engine may have a large fan diameter toprovide improved propulsive efficiency, whilst also having a relativelysmall diameter core. The may help to aid installation of the enginebeneath the wing of an aircraft. A relatively smaller diameter core mayallow the engine to be mounted relatively further aft on the aircraft,and so may allow the centre of gravity of the engine to be moved closerto the wing structure. The range of the fan tip axis angle defined abovemay allow improved engine mounting compared to known gas turbineengines, or where engine components are scaled proportionally withincreasing fan diameter. The fan axis angle may be higher than that ofknown gas turbine engines.

The fan axis angle may be in a range of 12 degrees and 17 degrees, forexample 13 degrees to 16 degrees.

The fan tip radius, measured as the radial distance from the enginecentreline to the radially outer tip of the leading edge of one of theplurality of fan blades, may be in a range from 95 cm to 200 cm, andoptionally may be from 110 cm to 150 cm or from 155 cm to 200 cm.

The fan axis angle may be in a range between 13 and 15 degrees—this maybe for an engine with a fan tip radius in the range from 110 cm to 150cm. The fan axis angle may be in a range between 13.5 degrees and 15.5degrees—this may be for an engine with a fan tip radius in the rangefrom 155 cm to 200 cm

The turbine radius at the lowest pressure rotor stage, measured/definedas the radial distance from the engine centreline to the radially outertip of the trailing edge of one of the rotor blades of the lowestpressure stage of the turbine, may be in the range from 45 cm to 85 cm.Optionally, for example for an engine with a fan tip radius in a rangefrom 110 cm to 150 cm, the turbine radius at the lowest pressure rotorstage may be in the range from 50 cm to 60 cm. Optionally, for examplefor an engine with a fan tip radius in a range from 155 cm to 200 cm,the turbine radius at the lowest pressure rotor stage may be in therange from range from 60 cm to 85 cm.

The gas turbine engine may further comprise a gearbox that receives aninput from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft.

The gas turbine engine may comprise more than one turbine. The turbinemay be a first turbine, the compressor may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft are arranged to rotateat a higher rotational speed than the first core shaft.

According to another aspect, there is provided a gas turbine engine foran aircraft comprising: an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor;and a fan located upstream of the engine core, the fan comprising aplurality of fan blades extending radially from a hub, each fan bladehaving a leading edge and a trailing edge; and a gearbox (30) thatreceives an input from the core shaft (26) and outputs drive to the fan(23) so as to drive the fan (23) at a lower rotational speed than thecore shaft (26); and wherein: the turbine comprises a lowest pressureturbine stage having a row of rotor blades, each of the rotor bladesextending radially and having a leading edge and a trailing edge; afan-turbine radius difference is measured as the radial distancebetween: a point on a circle swept by a radially outer tip of thetrailing edge of each of the rotor blades of the lowest pressure stageof the turbine; and a point on a circle swept by a radially outer tip ofthe leading edge of each of fan blades, and a fan speed to fan-turbineradius ratio defined as:

$\frac{{the}\mspace{14mu}{maximum}\mspace{14mu}{take}\text{-}{off}\mspace{14mu}{rotational}\mspace{14mu}{speed}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}\left( {{in}\mspace{14mu}{rpm}} \right)}{{fan}\text{-}{turbine}\mspace{14mu}{radius}\mspace{14mu}{difference}\mspace{14mu}\left( {{in}\mspace{14mu}{mm}} \right)}$is in a range of 0.8 rpm/mm to 5 rpm/mm.

By configuring the geometry of the gas turbine engine so that thefan-turbine radius difference is within the range above the loading onthe pylon which connects the gas turbine engine to the wing of anaircraft may be reduced. By defining the fan speed to fan-turbine radiusdifference ratio in this way a reduction in fan rotational speed mayreduce the restraining torque required for the pylon to restrain the gasturbine engine for relative rotation. Moreover, defining the ratio inthis range may also provide a smaller relative core diameter, allowingthe gas turbine engine to be mounted closer to the wing. This may alsoreduce the restraining torque on the pylon. The fan speed to fan-turbineradius may be lower than that of known gas turbine engines.

The fan speed to fan-turbine radius ratio may be in a range between 1.5rpm/mm to 4.0 rpm/mm. The fan speed to fan-turbine radius ratio may bein a range between 1.5 rpm/mm to 3.6 rpm/mm. The fan speed tofan-turbine radius ratio may be in a range between 2.9 rpm/mm and 3.8rpm/mm, for example 3.4 rpm/mm and 3.6 rpm/mm—this may be for an enginewith a fan tip radius in the range from 110 cm to 150 cm. The fan speedto fan-turbine radius ratio may be in a range between 1.2 rpm/mm and 2rpm/mm, for example 1.5 rpm/mm and 1.7 rpm/mm—this may be for an enginewith a fan tip radius in the range from 155 cm to 200 cm.

The fan-turbine radius difference may be in a range between 55 cm (i.e.550 mm) to 120 cm (i.e. 1200 mm). Optionally, for example for an enginewith a fan tip radius in the range from 110 cm to 150 cm, thefan-turbine radius difference may be in a range between 55 cm (i.e. 550mm) to 90 cm (i.e. 900 mm), for example 65 cm to 85 cm. Optionally, forexample for an engine with a fan tip radius in the range from 155 cm to200 cm, the fan-turbine radius difference may be in a range between 90cm (i.e. 900 mm) to 120 cm (i.e. 1200 mm), for example 95 cm to 115 cm.

The maximum take-off rotational fan speed may be in a range between 1450rpm to 3020 rpm. Optionally, for example for an engine with a fan tipradius in the range from 110 cm to 150 cm, the maximum take-offrotational fan speed may be in a range between 2100 rpm to 3000 rpm,optionally 2300 rpm to 2900 rpm. Optionally, for example for an enginewith a fan tip radius in the range from 155 cm to 200 cm, the maximumtake-off rotational fan speed may be in a range between 1450 rpm to 1910rpm, optionally 1500 rpm to 1800 rpm.

In embodiments in which the rotor (of the lowest pressure rotor stage)is shrouded, the radially outer tip of the trailing edge of each of therotor blades may be defined at the underside of the shroud. Inembodiments in which the rotor (of the lowest pressure rotor stage) isunshrouded, the radially outer tip of the trailing edge of each of therotor blades may be the blade tip of the rotor blade.

The turbine may be the lowest pressure turbine of a plurality ofturbines provided in the core. The turbine stage may be the axially mostrearward turbine stage and/or the most downstream turbine stage.

The gearbox may have a gear ratio in the range of from 3.2 to 5, forexample 3.2 to 4.2, for example 3.2 to 3.8.

The gas turbine engine of this aspect may further have a fan axis angleas defined in the previous relevant aspect or any of the abovestatements.

In another aspect, the present application provides a method ofoperating an aircraft comprising a gas turbine engine comprising: anengine core comprising a turbine, a compressor, and a core shaftconnecting the turbine to the compressor; and a fan located upstream ofthe engine core, the fan comprising a plurality of fan blades extendingradially from a hub, each fan blade having a leading edge and a trailingedge; and wherein: the turbine comprises a lowest pressure turbine stagehaving a row of rotor blades, each of the rotor blades extendingradially and having a leading edge and a trailing edge; a fan-turbineradius difference is measured as the radial distance between: a point ona circle swept by a radially outer tip of the trailing edge of each ofthe rotor blades of the lowest pressure stage of the turbine; and apoint on a circle swept by a radially outer tip of the leading edge ofeach of fan blades, wherein the method comprises controlling theaircraft such that a fan speed to fan-turbine radius ratio is definedas:

$\frac{{the}\mspace{14mu}{maximum}\mspace{14mu}{take}\text{-}{off}\mspace{14mu}{rotational}\mspace{14mu}{speed}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{fan}}{{fan}\text{-}{turbine}\mspace{14mu}{radius}\mspace{14mu}{difference}}$is in a range between 0.8 rpm/mm to 5 rpm/mm.

The gas turbine engine of this aspect may further have a fan axis angleas defined in the previous relevant aspect or any of the abovestatements

According to another aspect there is provided a gas turbine engine foran aircraft, and arranged to be mounted beneath a wing of the aircraft,the engine comprising:

an engine core comprising a turbine, a compressor, and a core shaftconnecting the turbine to the compressor, the turbine comprising alowest pressure rotor stage, the turbine having a turbine diameter;

a fan located upstream of the engine core, the fan comprising aplurality of fan blades extending from a hub; and

a gearbox that receives an input from the core shaft and outputs driveto the fan so as to drive the fan at a lower rotational speed than thecore shaft. A downstream blockage ratio of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{turbine}\mspace{14mu}{diameter}\mspace{14mu}{at}\mspace{14mu}{an}\mspace{14mu}{axial}\mspace{14mu}{location}} \\{{{of}\mspace{14mu}{the}\mspace{11mu}{lowest}\mspace{11mu}{pressure}\mspace{14mu}{rotor}\mspace{14mu}{stage}}\;}\end{matrix}}{{ground}\mspace{14mu}{plane}\mspace{14mu}{to}\mspace{14mu}{wing}\mspace{14mu}{distance}}$

is in the range from 0.2 to 0.3.

The present aspect relates to a gas turbine arranged to provide aspecified downstream blockage when mounted on the wing of an aircraft,the downstream blockage providing a measure of how much of the verticalspace beneath a wing of an aircraft is taken up by the gas turbineengine. The skilled person would appreciate that having a relatively lowturbine diameter towards the rearward end of the engine, as compared todistance from the aircraft wing on which the engine is to be mounted tothe ground plane, may allow more space for a pylon structure arranged tomount the engine to the wing. A lower downstream blockage may thereforebe preferable. In general, aircraft engines are mounted under aircraftwings, by means of a pylon (also referred to as an airframe strut)connected between a rearward portion of the engine and an underside ofthe wing. The turbine diameter at the axial position of the lowestpressure rotor may be used as a measure of engine size towards the rearof the engine.

The skilled person would appreciate that the ground plane is the planeon which the aircraft would land—i.e. the plane of the ground when theaircraft has landed/before take-off/whilst taxiing. The distance betweenthe ground plane and the wing may therefore be measured between theground (e.g. a runway) and the wing whilst the aircraft is parked. Theskilled person would appreciate that the ground plane is generally theplane in which the lowest part of each landing gear is located whilstthe aircraft is on the ground. The skilled person would appreciate thatthe distance between the ground plane and the wing may be measured tothe leading edge centre line of the aircraft wing.

The downstream blockage ratio may be lower than that of known aircraftgas turbine engines. The downstream blockage ratio may be in the rangefrom 0.20 to 0.30. The downstream blockage ratio may be in the rangefrom 0.20 to 0.29. The downstream blockage ratio may be in the rangefrom 0.20 to 0.28, and optionally from 0.22 to 0.28. Optionally, forexample for an engine with a fan tip radius in the range from 110 cm to150 cm, the downstream blockage ratio may be in the range from 0.23 to0.27. Optionally, for example for an engine with a fan tip radius in therange from 155 cm to 200 cm, downstream blockage ratio may be in therange from 0.24 to 0.28.

The distance between the ground plane and the wing may be measured tocentre point of a leading edge of the wing. The distance between theground plane and the wing may be measured along a line perpendicular tothe ground plane and passing through and perpendicular to an axialcentre line of the engine.

The turbine diameter at the axial location of the lowest pressure rotorstage may be measured adjacent blade tips of rotor blades of the lowestpressure rotor stage. The turbine diameter at the axial location of thelowest pressure rotor stage may be in the range from 70 cm to 170 cm.Optionally, for example for an engine with a fan tip radius in the rangefrom 110 cm to 150 cm, the turbine diameter at the lowest pressure rotorstage may be in the range from 100 cm to 130 cm, optionally 110 cm to120 cm. Optionally, for example for an engine with a fan tip radius inthe range from 155 cm to 200 cm, the turbine diameter at the lowestpressure rotor stage may be in the range from 120 cm to 170 cm.

The hub and fan blades of the fan together define a fan face having afan diameter, and the engine has an engine length. An engine ratio maybe defined as:

$\frac{\left( {2 \times {the}\mspace{14mu}{fan}\mspace{14mu}{{radius}/{the}}\mspace{14mu}{engine}\mspace{14mu}{length}} \right)}{{the}\mspace{14mu}{downstream}\mspace{14mu}{blockage}\mspace{14mu}{ratio}}$

The engine ratio may be in the range from 2.5 to 4. The engine ratio maybe in the range from 2.5 to 4.0. The engine ratio may be greater than2.5. The engine ratio may be greater than 3.0. The engine ratio may bein the range of from 2.7 to 3.7, for example 2.8 to 3.5.

The engine length may be measured as the axial distance between aforward region of the fan and a rearward region of the turbine. Theengine length may be measured along a centreline of the engine from anaxial position of an intersection of the leading edge of each fan bladeand the hub of the fan to an axial position of a trailing edge meanradius point of one of the rotor blades provided in the lowest pressurestage of the turbine.

The engine length may be in the range from 200 cm to 500 cm, andoptionally from 230 cm to 470 cm, optionally 300 cm to 450 cm.Optionally, for example for an engine with a fan tip radius in the rangefrom 110 cm to 150 cm (for example 120 cm to 140 cm), the engine lengthmay be in the range from 230 cm to 370 cm, optionally 300 to 360 cm.Optionally, for example for an engine with a fan tip radius in the rangefrom 155 cm to 200 cm (for example 165 cm to 190 cm), the engine lengthmay be in the range from 370 cm to 470 cm, optionally 390 cm to 450 cm.

The fan radius (also referred to as a fan tip radius) may be measuredbetween an engine centreline and a tip of a fan blade at its leadingedge and the fan diameter may be twice the radius of the fan. Thediameter of the fan, equal to twice the radius of the fan may be greaterthan (or on the order of) any of: 220 cm, 230 cm, 240 cm, 250 cm, 260cm, 270 cm, 280 cm, 290 cm, 300 cm, 310 cm, 320 cm, 330 cm, 340 cm, 350cm, 360 cm, 370 cm, 380 cm or 390 cm. In some embodiments, the fandiameter may be in the range from 220 cm to 300 cm. In some embodiments,the fan diameter may be in the range from 310 cm to 400 cm.

The turbine diameter at the lowest pressure rotor stage may be measuredat the axial location of blade tip trailing edges of rotor blades of thelowest pressure rotor stage. In embodiments in which the rotor (of thelowest pressure rotor stage) is shrouded, the turbine diameter of theturbine at the lowest pressure rotor stage may be measured to theunderside of the shroud (as this defines the edge of the gas flow path).In embodiments in which the rotor (of the lowest pressure rotor stage)is unshrouded, the turbine diameter of the turbine at the lowestpressure rotor stage may be measured to the blade tips of the rotor.

The gas turbine engine may comprise more than one turbine. The turbinemay be a first turbine, the compressor may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

A quasi-non-dimensional mass flow rate, Q, may be defined as:

$Q = {W\frac{\sqrt{T\; 0}}{P\;{0 \cdot A_{flow}}}}$where:W is mass flow rate through the fan in Kg/s;T0 is average stagnation temperature of the air at the fan face inKelvin;P0 is average stagnation pressure of the air at the fan face in Pa; andA_(flow) is the flow area of the fan face in m².A Q ratio of:

-   -   the downstream blockage ratio×quasi non dimensional mass flow        rate Q        may be in a range from 0.005 Kgs⁻¹N⁻¹K^(1/2) to 0.011        Kgs⁻¹N⁻¹K^(1/2).

The Q ratio may be in a range from 0.0050 Kgs⁻¹N⁻¹K^(1/2) to 0.0110Kgs⁻¹N⁻¹K^(1/2), or optionally from 0.005 Kgs⁻¹N⁻¹K^(1/2) to 0.010Kgs⁻¹N⁻¹K^(1/2). The Q ratio may be in a range from 0.006Kgs⁻¹N⁻¹K^(1/2) to 0.009 Kgs⁻¹N⁻¹K^(1/2).

A specific thrust may be defined as net engine thrust divided by massflow rate through the engine; and at engine cruise conditions, it may bethat:0.029Kgs ⁻¹ N ⁻¹ K ^(1/2) ≤Q≤0.036Kgs ⁻¹ N ⁻¹ K ^(1/2); and70Nkg ⁻¹ s≤specific thrust≤110Nkg ⁻¹ s.

At cruise conditions, it may be that: 0.032 Kgs⁻¹N⁻¹K^(1/2)≤Q≤0.036Kgs⁻¹N⁻¹K^(1/2). At cruise conditions, it may be that: 0.033Kgs⁻¹N⁻¹K^(1/2)≤Q≤0.035 Kgs⁻¹N⁻¹K^(1/2).

According to another aspect, there is provided an aircraft comprising awing and a gas turbine engine mounted beneath the wing of the aircraft.The engine comprises:

an engine comprising a turbine, a compressor, and a core shaftconnecting the turbine to the compressor, the turbine comprising aplurality of rotor stages including a lowest pressure rotor locatedfurthest downstream, the turbine having a turbine diameter;

a fan located upstream of the engine core, the fan comprising aplurality of fan blades; and

a gearbox that receives an input from the core shaft and outputs driveto the fan so as to drive the fan at a lower rotational speed than thecore shaft. A downstream blockage ratio of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{turbine}\mspace{14mu}{diameter}\mspace{14mu}{at}\mspace{14mu}{an}\mspace{14mu}{axial}\mspace{14mu}{location}\mspace{14mu}{of}\mspace{14mu}{the}} \\{{lowest}\mspace{14mu}{pressure}\mspace{14mu}{rotor}}\end{matrix}}{{ground}\mspace{14mu}{plane}\mspace{14mu}{to}\mspace{14mu}{wing}\mspace{14mu}{distance}}$

is in the range from 0.2 to 0.3.

The engine may be as described for the preceding aspect.

According to another aspect, there is provided a gas turbine engine foran aircraft and arranged to be mounted beneath a wing of the aircraft.The engine has an engine length and comprises:

an engine core comprising a turbine, a compressor, and a core shaftconnecting the turbine to the compressor, the turbine comprising alowest pressure rotor stage, the turbine having a turbine diameter;

a fan located upstream of the engine core, the fan comprising aplurality of fan blades extending from a hub, the hub and fan blades ofthe fan together defining a fan face having a fan tip radius; and

a gearbox that receives an input from the core shaft and outputs driveto the fan so as to drive the fan at a lower rotational speed than thecore shaft. A downstream blockage ratio is defined as:

$\frac{\begin{matrix}{{the}\mspace{14mu}{turbine}\mspace{14mu}{diameter}\mspace{14mu}{at}\mspace{14mu}{an}\mspace{14mu}{axial}\mspace{14mu}{location}\mspace{14mu}{of}\mspace{14mu}{the}} \\{{lowest}\mspace{14mu}{pressure}\mspace{14mu}{rotor}}\end{matrix}}{{ground}\mspace{14mu}{plane}\mspace{14mu}{to}\mspace{14mu}{wing}\mspace{14mu}{distance}}$An engine blockage ratio of:

$\frac{\left( {2 \times {the}\mspace{14mu}{fan}{\mspace{11mu}\;}{tip}\mspace{14mu}{{radius}/{the}}\mspace{14mu}{engine}\mspace{14mu}{length}} \right)}{{the}\mspace{14mu}{downstream}\mspace{14mu}{blockage}\mspace{14mu}{ratio}}$

is in the range from 2.5 to 4.

The engine blockage ratio may equivalently be represented as:

$\frac{2 \times {the}\mspace{14mu}{fan}\mspace{14mu}{tip}\mspace{14mu}{radius} \times {ground}\mspace{14mu}{plane}\mspace{14mu}{to}\mspace{14mu}{wing}\mspace{14mu}{distance}}{\begin{matrix}{{engine}\mspace{14mu}{length} \times {turbine}\mspace{14mu}{diameter}\mspace{14mu}{at}\mspace{14mu}{an}\mspace{14mu}{axial}\mspace{14mu}{location}} \\{{of}\mspace{14mu}{the}\mspace{14mu}{lowest}\mspace{14mu}{pressure}\mspace{14mu}{rotor}}\end{matrix}}$

The skilled person would appreciate that a larger fan diameter (andtherefore a larger fan tip radius) may improve propulsive efficiency,but that simply mounting a larger fan on a known gas turbine engine maycause other difficulties or inefficiencies. Having a shorter enginelength may allow an engine with a larger fan to be installed closer to awing of the aircraft, i.e. further rearwards; this may reduce the momentexerted on the wing by the (potentially larger and heavier) engine.Similarly, reducing the turbine diameter may allow an engine with alarger fan to be installed higher up (maintaining intake groundclearance) whilst maintaining adequate vertical depth for a pylonstructure arranged to connect the engine to the wing. The engineblockage ratio may be higher than that of known aircraft gas turbineengines.

The engine blockage ratio may be in the range from 2.5 to 4.0,optionally 2.7 to 3.7. The engine blockage ratio may be greater than2.5. The engine blockage ratio may be greater than 3.0.

The downstream blockage ratio may be as described in the precedingaspect. The downstream blockage ratio may be in the range from 0.2 to0.3. The downstream blockage ratio may be in the range from 0.20 to0.30. The downstream blockage ratio may be in the range from 0.20 to0.29. The downstream blockage ratio may be in the range from 0.22 to0.28.

The distance between the ground plane and the wing may be measured tocentre point of a leading edge of the wing. The distance between theground plane and the wing may be measured along a line perpendicular tothe ground plane and passing through and perpendicular to an axialcentre line of the engine.

The turbine diameter at the axial location of the lowest pressure rotorstage may be measured adjacent blade tips of rotor blades of the lowestpressure rotor stage. The turbine diameter at the axial location of thelowest pressure rotor stage may be in the range from 70 cm to 170 cm.Optionally, for example for an engine with a fan tip radius in the rangefrom 110 cm to 150 cm, the turbine diameter at the lowest pressure rotorstage may be in the range from 100 cm to 130 cm, optionally 110 cm to120 cm. Optionally, for example for an engine with a fan tip radius inthe range from 155 cm to 200 cm, the turbine diameter at the lowestpressure rotor stage may be in the range from 120 cm to 170 cm.

The engine length may be measured as the axial distance between aforward region of the fan and a rearward region of the turbine. Theengine length may be measured along a centreline of the engine from anaxial position of an intersection of the leading edge of each fan bladeand the hub of the fan to an axial position of a trailing edge meanradius point of one of the rotor blades provided in the lowest pressurestage of the turbine.

The engine length may be in the range from 200 cm to 500 cm, andoptionally from 230 cm to 470 cm, optionally 300 cm to 450 cm.Optionally, for example for an engine with a fan tip radius in the rangefrom 110 cm to 150 cm (for example 120 cm to 140 cm), the engine lengthmay be in the range from 230 cm to 370 cm, optionally 300 to 360 cm.Optionally, for example for an engine with a fan tip radius in the rangefrom 155 cm to 200 cm (for example 165 cm to 190 cm), the engine lengthmay be in the range from 370 cm to 470 cm, optionally 390 cm to 450 cm.

The fan tip radius may be measured between an engine centreline and atip of a fan blade at its leading edge. A fan diameter equal to twicethe fan tip radius may be greater than (or on the order of) any of: 220cm, 230 cm, 240 cm, 250 cm, 260 cm, 270 cm, 280 cm, 290 cm, 300 cm, 310cm, 320 cm, 330 cm, 340 cm, 350 cm, 360 cm, 370 cm, 380 cm or 390 cm.

The turbine may be a first turbine, the compressor may be a firstcompressor, and the core shaft may be a first core shaft. The enginecore may further comprise a second turbine, a second compressor, and asecond core shaft connecting the second turbine to the secondcompressor. The second turbine, second compressor, and second core shaftmay be arranged to rotate at a higher rotational speed than the firstcore shaft.

A quasi-non-dimensional mass flow rate, Q, may be defined as:

$Q = {W\frac{\sqrt{T\; 0}}{P\;{0 \cdot A_{flow}}}}$where:W is mass flow rate through the fan in Kg/s;T0 is average stagnation temperature of the air at the fan face inKelvin;P0 is average stagnation pressure of the air at the fan face in Pa; andA_(flow) is the flow area of the fan face in m².A Q ratio of:

-   -   the downstream blockage ratio×quasi non dimensional mass flow        rate Q        may be in a range from 0.005 Kgs⁻¹N⁻¹K^(1/2) to 0.011        Kgs⁻¹N⁻¹K^(1/2).

The Q ratio may be in a range from 0.005 Kgs⁻¹N⁻¹K^(1/2) to 0.010Kgs⁻¹N⁻¹K^(1/2) from 0.0050 Kgs⁻¹N⁻¹K^(1/2) to 0.0110 Kgs⁻¹N⁻¹K^(1/2),or optionally from 0.0050 Kgs⁻¹N⁻¹K^(1/2) to 0.0100 Kgs⁻¹N⁻¹K^(1/2). TheQ ratio may be in a range from 0.006 Kgs⁻¹N⁻¹K^(1/2) to 0.009Kgs⁻¹N⁻¹K^(1/2).

A specific thrust may be defined as net engine thrust divided by massflow rate through the engine; and at engine cruise conditions, it may bethat:0.029Kgs ⁻¹ N ⁻¹ K ^(1/2) ≤Q≤0.036Kgs ⁻¹ N ⁻¹ K ^(1/2); and70Nkg ⁻¹ s≤specific thrust≤110Nkg ⁻¹ s.

At cruise conditions, it may be that: 0.032 Kgs⁻¹N⁻¹K^(1/2)≤Q≤0.036Kgs⁻¹N⁻¹K^(1/2). At cruise conditions, it may be that: 0.033Kgs⁻¹N⁻¹K^(1/2)≤Q≤0.035 Kgs⁻¹N⁻¹K^(1/2).

According to another aspect, there is provided an aircraft comprising awing and a gas turbine engine mounted beneath the wing of the aircraft.The engine has an engine length and comprises:

an engine core comprising a turbine, a compressor, and a core shaftconnecting the turbine to the compressor, the turbine comprising alowest pressure rotor stage, the turbine having a turbine diameter;

a fan located upstream of the engine core, the fan comprising aplurality of fan blades extending from a hub, the hub and fan blades ofthe fan together defining a fan face having a fan tip radius; and

a gearbox that receives an input from the core shaft and outputs driveto the fan so as to drive the fan at a lower rotational speed than thecore shaft. A downstream blockage ratio is defined as:

$\frac{\begin{matrix}{{the}\mspace{14mu}{turbine}\mspace{14mu}{diameter}\mspace{14mu}{at}\mspace{14mu}{an}\mspace{14mu}{axial}\mspace{14mu}{location}\mspace{14mu}{of}\mspace{14mu}{the}} \\{{lowest}\mspace{14mu}{pressure}\mspace{14mu}{rotor}}\end{matrix}}{{ground}\mspace{14mu}{plane}\mspace{14mu}{to}\mspace{14mu}{wing}\mspace{14mu}{distance}}$An engine blockage ratio of:

$\frac{\left( {2 \times {the}\mspace{14mu}{fan}{\mspace{11mu}\;}{tip}\mspace{14mu}{{radius}/{the}}\mspace{14mu}{engine}\mspace{14mu}{length}} \right)}{{the}\mspace{14mu}{downstream}\mspace{14mu}{blockage}\mspace{14mu}{ratio}}$is in the range from 2.5 to 4.

The downstream blockage ratio may be in the range from 0.2 to 0.3. Theengine may be as described for the preceding aspect.

In another aspect, there is provided a gas turbine engine for anaircraft and arranged to be mounted beneath a wing of the aircraft, thegas turbine engine comprising:

an engine core comprising a turbine, a compressor, and a core shaftconnecting the turbine to the compressor, the turbine comprising alowest pressure rotor stage, the turbine having a turbine diameter;

a fan located upstream of the engine core, the fan comprising aplurality of fan blades extending from a hub, an annular fan face beingdefined at a leading edge of the fan; and wherein:

a downstream blockage ratio is defined as the ratio of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{turbine}\mspace{14mu}{diameter}\mspace{14mu}{at}\mspace{14mu}{an}\mspace{14mu}{axial}\mspace{14mu}{location}\mspace{14mu}{of}\mspace{14mu}{the}} \\{{lowest}\mspace{14mu}{pressure}\mspace{14mu}{rotor}\mspace{14mu}{stage}}\end{matrix}}{{ground}\mspace{14mu}{plane}\mspace{14mu}{to}\mspace{14mu}{wing}\mspace{14mu}{distance}}$and a quasi-non-dimensional mass flow rate Q is defined as:

$Q = {W\frac{\sqrt{T\; 0}}{P\;{0 \cdot A_{flow}}}}$

where:

W is mass flow rate through the fan in Kg/s;

T0 is average stagnation temperature of the air at the fan face inKelvin;

P0 is average stagnation pressure of the air at the fan face in Pa; and

A_(flow) is the flow area of the fan face in m²

and wherein a Q ratio of:

-   -   the downstream blockage ratio×Q        is in a range from 0.005 Kgs⁻¹N⁻¹K^(1/2) to 0.011        Kgs⁻¹N⁻¹K^(1/2). The values of Q used to calculate the Q ratio        as referred to herein may be at cruise conditions.

By defining the Q ratio in this range a large mass flow may be achievedwhile also minimising the downstream blockage. The Q ratio may be lowerthan that of a known gas turbine engine. The Q ratio may be in a rangefrom 0.0050 Kgs⁻¹N⁻¹K^(1/2) to 0.0110, or to 0.0100 Kgs⁻¹N⁻¹K^(1/2). TheQ ratio may be in a range from 0.005 Kgs⁻¹N⁻¹K^(1/2) to 0.010Kgs⁻¹N⁻¹K^(1/2). The Q ratio may be in a range from 0.006Kgs⁻¹N⁻¹K^(1/2) to 0.009 Kgs⁻¹N⁻¹K^(1/2).

The downstream blockage ratio may be in a range from 0.2 to 0.3. Thedownstream blockage ratio may be in a range from 0.20 to 0.29. Thedownstream blockage ratio may be in a range from 0.22 to 0.28.

The distance between the ground plane and the wing may be measured to acentre point of a leading edge of the wing. The distance between theground plane and the wing may be measured along a line perpendicular tothe ground plane and passing through and perpendicular to an axialcentreline of the engine.

The turbine diameter at the axial location of the lowest pressure rotorstage may be measured adjacent blade tips of rotor blades of the lowestpressure rotor stage. The turbine diameter at the axial location of thelowest pressure rotor stage may be in the range from 70 cm to 170 cm.Optionally, for example for an engine with a fan tip radius in the rangefrom 110 cm to 150 cm, the turbine diameter at the lowest pressure rotorstage may be in the range from 100 cm to 120 cm. Optionally, for examplefor an engine with a fan tip radius in the range from 155 cm to 200 cm,the turbine diameter at the lowest pressure rotor stage may be in therange from 120 cm to 170 cm. The turbine diameter at the lowest pressurerotor stage may be measured at the axial location of blade tip trailingedges of rotor blades of the lowest pressure rotor stage. The turbinediameter of the turbine at the lowest pressure rotor stage may bemeasured:

-   -   (i) when the rotor is shrouded, to the underside of the shroud;        or    -   (ii) when the rotor is unshrouded, to the blade tips of the        rotor.

A specific thrust may be defined as net engine thrust divided by massflow rate through the engine; and at engine cruise conditions it may bethat:0.029Kgs ⁻¹ N ⁻¹ K ^(1/2) ≤Q≤0.036Kgs ⁻¹ N ⁻¹ K ^(1/2); and/or70Nkg ⁻¹ s≤specific thrust≤110Nkg ⁻¹ s.

It may be that at cruise conditions: 0.032 Kgs⁻¹N⁻¹K^(1/2)≤Q≤0.036Kgs⁻¹N⁻¹K^(1/2). It may be that at cruise conditions: 0.033Kgs⁻¹N⁻¹K^(1/2)≤Q≤0.035 Kgs⁻¹N⁻¹K^(1/2).

The ratio of the radius of fan blade at its hub to the radius of the fanblade at its tip may be less than 0.33.

At cruise conditions, the specific thrust may be less than 100 Nkg⁻¹s.The forward speed of the gas turbine engine at the cruise conditions maybe in the range of from Mn 0.75 to Mn 0.85. The forward speed of the gasturbine engine at the cruise conditions may be Mn 0.8. The cruiseconditions may correspond to atmospheric conditions at an altitude thatis in the range of from 10500 m to 11600 m. The cruise conditions maycorrespond to atmospheric conditions at an altitude of 11000 m. Thecruise conditions may correspond to:

a forward Mach number of 0.8;

a pressure of 23000 Pa; and

a temperature of −55 deg C.

The gas turbine engine may comprise a gearbox that receives an inputfrom the core shaft and outputs drive to the fan so as to drive the fanat a lower rotational speed than the core shaft.

The turbine may be a first turbine, the compressor may be a firstcompressor, and the core shaft may be a first core shaft. The enginecore may further comprise a second turbine, a second compressor, and asecond core shaft connecting the second turbine to the secondcompressor. The second turbine, second compressor, and second core shaftmay be arranged to rotate at a higher rotational speed than the firstcore shaft.

In another aspect, there is provided an aircraft comprising a wing and agas turbine engine mounted beneath the wing of the aircraft, the enginecomprising: an engine core comprising a turbine, a compressor, and acore shaft connecting the turbine to the compressor, the turbinecomprising a lowest pressure rotor stage, the turbine having a turbinediameter; and a fan located upstream of the engine core, the fancomprising a plurality of fan blades extending from a hub, an annularfan face being defined at a leading edge of the fan wherein: adownstream blockage ratio is defined as the ratio of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{turbine}\mspace{14mu}{diameter}\mspace{14mu}{at}\mspace{14mu}{an}\mspace{14mu}{axial}\mspace{14mu}{location}\mspace{14mu}{of}\mspace{14mu}{the}} \\{{lowest}\mspace{14mu}{pressure}\mspace{14mu}{rotor}\mspace{14mu}{stage}}\end{matrix}}{{ground}\mspace{14mu}{plane}\mspace{14mu}{to}\mspace{14mu}{wing}\mspace{14mu}{distance}}$and a quasi-non-dimensional mass flow rate Q is defined as:

$Q = {W\frac{\sqrt{T\; 0}}{P\;{0 \cdot A_{flow}}}}$where:W is mass flow rate through the fan in Kg/s;T0 is average stagnation temperature of the air at the fan face inKelvin;P0 is average stagnation pressure of the air at the fan face in Pa; andA_(flow) is the flow area of the fan face in m²and wherein a Q ratio of:

-   -   the downstream blockage ratio×Q        is in a range from 0.005 Kgs⁻¹N⁻¹K^(1/2) to 0.011        Kgs⁻¹N⁻¹K^(1/2).        The aircraft may comprise two wings, with one or more gas        turbine engines mounted beneath each wing; the or each gas        turbine engine may have the features of any of the statements        related to the previous aspect.        According to another aspect, there is provided a method of        operating an aircraft comprising a gas turbine engine        comprising: an engine core comprising a turbine, a compressor,        and a core shaft connecting the turbine to the compressor, the        turbine comprising a lowest pressure rotor stage, the turbine        having a turbine diameter; and a fan located upstream of the        engine core, the fan comprising a plurality of fan blades        extending from a hub, an annular fan face being defined at a        leading edge of the fan, wherein: a downstream blockage ratio is        defined as the ratio of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{turbine}\mspace{14mu}{diameter}\mspace{14mu}{at}\mspace{14mu}{an}\mspace{14mu}{axial}\mspace{14mu}{location}\mspace{14mu}{of}\mspace{14mu}{the}} \\{{lowest}\mspace{14mu}{pressure}\mspace{14mu}{rotor}\mspace{14mu}{stage}}\end{matrix}}{{ground}\mspace{14mu}{plane}\mspace{14mu}{to}\mspace{14mu}{wing}\mspace{14mu}{distance}}$and a quasi-non-dimensional mass flow rate Q is defined as:

$Q = {W\frac{\sqrt{T\; 0}}{P\;{0 \cdot A_{flow}}}}$where:W is mass flow rate through the fan in Kg/s;T0 is average stagnation temperature of the air at the fan face inKelvin;P0 is average stagnation pressure of the air at the fan face in Pa; andA_(flow) is the flow area of the fan face in m²the method comprises controlling the aircraft such that a Q ratio of:the downstream blockage ratio×Q is in a range from 0.005 Kgs⁻¹N⁻¹K^(1/2)to 0.011 Kgs⁻¹N⁻¹K^(1/2).

The gas turbine engine may have the features of any of the statementsrelated to the previous two aspects.

The skilled person will appreciate that a feature described above inrelation to any one of the aspects may be applied, mutatis mutandis, toany other aspect of the invention. For example, in various embodimentsany two or more of the conditions for ratios as defined above, andoptionally all specified ratio ranges, may apply to any given aspect orembodiment. All aspects may apply to an engine of some embodiments.Furthermore, any feature described below may apply to any aspect and/ormay apply in combination with any one of the claims.

Arrangements of the present disclosure may be particularly, although notexclusively, beneficial for fans that are driven via a gearbox.Accordingly, the gas turbine engine may comprise a gearbox that receivesan input from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft. The input tothe gearbox may be directly from the core shaft, or indirectly from thecore shaft, for example via a spur shaft and/or gear. The core shaft mayrigidly connect the turbine and the compressor, such that the turbineand compressor rotate at the same speed (with the fan rotating at alower speed).

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example the first core shaft in the example above). For example,the gearbox may be arranged to be driven only by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example only be the first core shaft, and not the second coreshaft, in the example above). Alternatively, the gearbox may be arrangedto be driven by any one or more shafts, for example the first and/orsecond shafts in the example above.

The gearbox is a reduction gearbox (in that the output to the fan is alower rotational rate than the input from the core shaft). Any type ofgearbox may be used. For example, the gearbox may be a “planetary” or“star” gearbox, as described in more detail elsewhere herein. Thegearbox may have any desired reduction ratio (defined as the rotationalspeed of the input shaft divided by the rotational speed of the outputshaft), for example greater than 2.5, for example in the range of from 3to 5, for example on the order of or at least 3, 3.1, 3.2, 3.3, 3.4,3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratio may be, forexample, between any two of the values in the previous sentence. Thegear ratio may be in the range of from 3.2 to 4.2, and optionally in therange 3.2 to 3.8, 3.3 to 3.8 or 3.4 to 3.7. A higher gear ratio may bemore suited to “planetary” style gearbox. In some arrangements, the gearratio may be outside these ranges.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan andcompressor(s). For example, the combustor may be directly downstream of(for example at the exit of) the second compressor, where a secondcompressor is provided. By way of further example, the flow at the exitto the combustor may be provided to the inlet of the second turbine,where a second turbine is provided. The combustor may be providedupstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). These ratios may commonly be referred to as the hub-to-tipratio. The radius at the hub and the radius at the tip may both bemeasured at the leading edge (or axially forwardmost) part of the blade.The hub-to-tip ratio refers, of course, to the gas-washed portion of thefan blade, i.e. the portion radially outside any platform.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches),260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm(around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350cm, 360 cm (around 140 inches), 370 cm (around 145 inches), 380 (around150 inches) cm or 390 cm (around 155 inches). The fan diameter may be inan inclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds).

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 220 cm to 300 cm (for example 240 cm to 280 cm) may be inthe range of from 1700 rpm to 2500 rpm, for example in the range of from1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100rpm. Purely by way of further non-limitative example, the rotationalspeed of the fan at cruise conditions for an engine having a fandiameter in the range of from 320 cm to 380 cm may be in the range offrom 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to1800 rpm, for example in the range of from 1400 rpm to 1600 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades on the flow results in an enthalpy rise dH of the flow. A fan tiploading may be defined as dH/U_(tip) ², where dH is the enthalpy rise(for example the 1-D average enthalpy rise) across the fan and U_(tip)is the (translational) velocity of the fan tip, for example at theleading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.28, 0.29,0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (allunits in this paragraph being Jkg⁻¹K⁻¹/(ms⁻¹)²). The fan tip loading maybe in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds).

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5, or 20. The bypassratio may be in an inclusive range bounded by any two of the values inthe previous sentence (i.e. the values may form upper or lower bounds).The bypass ratio may be in the range of from 11 to 20, and optionally inthe range from 13 to 20 or 14 to 20.

The bypass duct may be substantially annular. The bypass duct may beradially outside the core engine. The radially outer surface of thebypass duct may be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressureupstream of the fan to the stagnation pressure at the exit of thehighest pressure compressor (before entry into the combustor). By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds).

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹s, 105 Nkg⁻¹s, 100 Nkg⁻¹s, 95 Nkg⁻¹s, 90 Nkg⁻¹s, 85 Nkg⁻¹s or 80Nkg⁻¹s. The specific thrust may be in an inclusive range bounded by anytwo of the values in the previous sentence (i.e. the values may formupper or lower bounds). Such engines may be particularly efficient incomparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). The thrust referred to abovemay be the maximum net thrust at standard atmospheric conditions at sealevel plus 15 deg C. (ambient pressure 101.3 kPa, temperature 30 degC.), with the engine static.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400K, 1450K, 1500K,1550K, 1600K or 1650K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). The maximum TET may occur, for example, at a high thrustcondition, for example at a maximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein according to any aspect may be manufactured from anysuitable material or combination of materials. For example at least apart of the fan blade and/or aerofoil may be manufactured at least inpart from a composite, for example a metal matrix composite and/or anorganic matrix composite, such as carbon fibre. By way of furtherexample at least a part of the fan blade and/or aerofoil may bemanufactured at least in part from a metal, such as a titanium basedmetal or an aluminium based material (such as an aluminium-lithiumalloy) or a steel based material. The fan blade may comprise at leasttwo regions manufactured using different materials. For example, the fanblade may have a protective leading edge, which may be manufacturedusing a material that is better able to resist impact (for example frombirds, ice or other material) than the rest of the blade. Such a leadingedge may, for example, be manufactured using titanium or atitanium-based alloy. Thus, purely by way of example, the fan blade mayhave a carbon-fibre or aluminium based body (such as an aluminiumlithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmay be formed integrally with a central portion. Such an arrangement maybe referred to as a blisk or a bling. Any suitable method may be used tomanufacture such a blisk or bling. For example, at least a part of thefan blades may be machined from a block and/or at least part of the fanblades may be attached to the hub/disc by welding, such as linearfriction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26fan blades.

As used herein, cruise conditions have the conventional meaning andwould be readily understood by the skilled person. Thus, for a given gasturbine engine for an aircraft, the skilled person would immediatelyrecognise cruise conditions to mean the operating point of the engine atmid-cruise of a given mission (which may be referred to in the industryas the “economic mission”) of an aircraft to which the gas turbineengine is designed to be attached. In this regard, mid-cruise is thepoint in an aircraft flight cycle at which 50% of the total fuel that isburned between top of climb and start of descent has been burned (whichmay be approximated by the midpoint—in terms of time and/ordistance—between top of climb and start of descent. Cruise conditionsthus define an operating point of the gas turbine engine that provides athrust that would ensure steady state operation (i e maintaining aconstant altitude and constant Mach Number) at mid-cruise of an aircraftto which it is designed to be attached, taking into account the numberof engines provided to that aircraft. For example where an engine isdesigned to be attached to an aircraft that has two engines of the sametype, at cruise conditions the engine provides half of the total thrustthat would be required for steady state operation of that aircraft atmid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruiseconditions are defined as the operating point of the engine thatprovides a specified thrust (required to provide—in combination with anyother engines on the aircraft—steady state operation of the aircraft towhich it is designed to be attached at a given mid-cruise Mach Number)at the mid-cruise atmospheric conditions (defined by the InternationalStandard Atmosphere according to ISO 2533 at the mid-cruise altitude).For any given gas turbine engine for an aircraft, the mid-cruise thrust,atmospheric conditions and Mach Number are known, and thus the operatingpoint of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be part of the cruise condition.For some aircraft, the cruise conditions may be outside these ranges,for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions (according to the International StandardAtmosphere, ISA) at an altitude that is in the range of from 10000 m to15000 m, for example in the range of from 10000 m to 12000 m, forexample in the range of from 10400 m to 11600 m (around 38000 ft), forexample in the range of from 10500 m to 11500 m, for example in therange of from 10600 m to 11400 m, for example in the range of from 10700m (around 35000 ft) to 11300 m, for example in the range of from 10800 mto 11200 m, for example in the range of from 10900 m to 11100 m, forexample on the order of 11000 m. The cruise conditions may correspond tostandard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to anoperating point of the engine that provides a known required thrustlevel (for example a value in the range of from 30 kN to 35 kN) at aforward Mach number of 0.8 and standard atmospheric conditions(according to the International Standard Atmosphere) at an altitude of38000 ft (11582 m). Purely by way of further example, the cruiseconditions may correspond to an operating point of the engine thatprovides a known required thrust level (for example a value in the rangeof from 50 kN to 65 kN) at a forward Mach number of 0.85 and standardatmospheric conditions (according to the International StandardAtmosphere) at an altitude of 35000 ft (10668 m).

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

According to an aspect, there is provided an aircraft comprising a gasturbine engine as described and/or claimed herein. The aircraftaccording to this aspect is the aircraft for which the gas turbineengine has been designed to be attached. Accordingly, the cruiseconditions and/or maximum take-off conditions according to this aspectmay correspond to those of the aircraft, as defined elsewhere herein.

According to an aspect, there is provided a method of operating a gasturbine engine as described and/or claimed herein. The operation may beat the cruise conditions and/or maximum take-off conditions as definedelsewhere herein (for example in terms of the thrust, atmosphericconditions and Mach Number).

According to an aspect, there is provided a method of operating anaircraft comprising a gas turbine engine as described and/or claimedherein. The operation according to this aspect may include (or may be)operation at the mid-cruise and/or maximum take-off conditions of theaircraft, as defined elsewhere herein.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close up sectional side view of an upstream portion of a gasturbine engine;

FIG. 3A is a partially cut-away view of a gearbox for a gas turbineengine;

FIG. 3B is a sectional view of the gas turbine engine of FIG. 1 withnozzle parameters marked;

FIG. 4A is a sectional view of the gas turbine engine of FIG. 1 withmarked engine dimensions suitable for use in calculating an engine arearatio;

FIG. 4B is a schematic sectional view of a generic gas turbine enginewith marked engine dimensions corresponding to those marked in FIG. 4A;

FIG. 4C is a schematic sectional view of a generic engine core and fanwith marked engine dimensions corresponding to those marked in FIG. 4A;

FIG. 5A is a sectional view of the gas turbine engine of FIG. 1 withmarked engine dimensions suitable for use in calculating an enginelength to CoG ratio;

FIG. 5B is a schematic sectional view of a generic gas turbine enginewith marked engine dimensions corresponding to those marked in FIG. 5A;

FIG. 5C illustrates a method of an embodiment;

FIG. 6A is a sectional view of the gas turbine engine of FIG. 1 withmarked engine dimensions suitable for use in calculating a gearboxlocation to engine length ratio;

FIG. 6B is a schematic sectional view of a generic gas turbine enginewith marked engine dimensions corresponding to those marked in FIG. 6A;

FIG. 7A is a sectional view of the gas turbine engine of FIG. 1 withmarked engine dimensions suitable for use in calculating an outer bypassto fan ratio;

FIG. 7B is a sectional view of a different gas turbine engine having adifferent nacelle shape, with marked engine dimensions suitable for usein calculating an outer bypass to fan ratio;

FIG. 7C is a schematic sectional view of a generic gas turbine enginewith marked engine dimensions corresponding to those marked in FIG. 7A;

FIG. 8A is a sectional view of the gas turbine engine of FIG. 1 withmarked engine dimensions suitable for use in calculating an inner bypassto fan ratio;

FIG. 8B is a schematic sectional view of a generic gas turbine enginewith marked engine dimensions corresponding to those marked in FIG. 8A;

FIG. 8C is a schematic sectional view of a generic gas turbine enginewith marked engine dimensions corresponding to those marked in FIG. 8A;

FIG. 9A is a sectional view of the gas turbine engine of FIG. 1 withmarked engine dimensions suitable for use in calculating a fan axisangle ratio;

FIG. 9B is a schematic sectional view of a generic gas turbine enginewith marked engine dimensions corresponding to those marked in FIG. 9A;

FIG. 10A is a sectional view of the gas turbine engine of FIG. 1 withmarked engine dimensions suitable for use in calculating a fan speed tofan-turbine radius difference ratio;

FIG. 10B is a schematic sectional view of a generic gas turbine enginewith marked engine dimensions corresponding to those marked in FIG. 10A;

FIG. 10C illustrates a method of an embodiment;

FIG. 11A is a sectional view of the gas turbine engine of FIG. 1 incontext between the ground and a wing of the aircraft, with markedengine dimensions suitable for use in calculating downstream blockage;

FIG. 11B is a schematic sectional view of a generic gas turbine enginewith marked engine dimensions corresponding to those marked in FIG. 11A;

FIG. 12A is a sectional view of the gas turbine engine of FIG. 1 withmarked engine dimensions suitable for use in calculating an outer bypassduct wall angle;

FIG. 12B is a schematic sectional view of a generic gas turbine enginewith marked engine dimensions corresponding to those marked in FIG. 12A;

FIG. 12C is a schematic sectional view of a generic gas turbine enginewith marked engine dimensions corresponding to another bypass duct wallangle;

FIG. 13A is a sectional view of the gas turbine engine of FIG. 1 withmarked engine dimensions suitable for use in calculating a bypass tocore ratio;

FIG. 13B is a schematic sectional view of a generic gas turbine enginewith marked engine dimensions corresponding to those marked in FIG. 13A;

FIG. 13C illustrates a method of an embodiment;

FIG. 14A is a schematic sectional view of the gas turbine engine of FIG.1 with marked engine dimensions suitable for use in calculating a quasinon-dimensional mass flow rate (Q);

FIG. 14B illustrates a method of an embodiment;

FIG. 15A is a schematic sectional view of an unshrouded turbine rotor ina radial plane;

FIG. 15B is a schematic sectional view of a shrouded turbine rotor in aradial plane; and

FIG. 16 is a schematic view of an aircraft comprising two gas turbineengines.

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, a lowpressure compressor 14, a high-pressure compressor 15, combustionequipment 16, a high-pressure turbine 17, a low pressure turbine 19 anda core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. Thebypass airflow B flows through the bypass duct 22. The fan 23 isattached to and driven by the low pressure turbine 19 via a shaft 26 andan epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel and the mixture is combusted.The resultant hot combustion products then expand through, and therebydrive, the high pressure and low pressure turbines 17, 19 before beingexhausted through the nozzle 20 to provide some propulsive thrust. Thehigh pressure turbine 17 drives the high pressure compressor 15 by asuitable interconnecting shaft 27. The fan 23 generally provides themajority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26,which is coupled to a sun wheel, or sun gear, 28 of the epicyclic geararrangement 30. Radially outwardly of the sun gear 28 and intermeshingtherewith is a plurality of planet gears 32 that are coupled together bya planet carrier 34. The planet carrier 34 constrains the planet gears32 to precess around the sun gear 28 in synchronicity whilst enablingeach planet gear 32 to rotate about its own axis. The planet carrier 34is coupled via linkages 36 to the fan 23 in order to drive its rotationabout the engine axis 9. Radially outwardly of the planet gears 32 andintermeshing therewith is an annulus or ring gear 38 that is coupled,via linkages 40, to a stationary supporting structure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3. There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIG. 2. For example, where the gearbox 30 has a stararrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2.

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine 10 shown in FIG. 1 has a split flow nozzle 18, 20meaning that the flow through the bypass duct 22 has its own nozzle 18that is separate to and radially outside the core engine nozzle 20.However, this is not limiting, and various aspects of the presentdisclosure may also apply to engines in which the flow through thebypass duct 22 and the flow through the core 11 are mixed, or combined,before (or upstream of) a single nozzle, which may be referred to as amixed flow nozzle. One or both nozzles (whether mixed or split flow) mayhave a fixed or variable area. Whilst the described example relates to aturbofan engine, the disclosure may apply, for example, to any type ofgas turbine engine, such as an open rotor (in which the fan stage is notsurrounded by a nacelle) or turboprop engine, for example. In somearrangements, the gas turbine engine 10 may not comprise a gearbox 30.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

Referring again to FIGS. 1 and 2, the lowest pressure compressor 14comprises one or more compressor stages. In the embodiment shown in FIG.1, the lowest pressure compressor 14 comprises two compressor stages.Each stage of the compressor may comprise a row of rotor blades 14 a, 14b and a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

The one or more compressor stages may comprise a lowest pressure stage,and may further comprise one or more compressor stages of increasingpressure to a highest pressure compressor stage. The lowest pressurecompressor stage 14 a may be located furthest upstream along the gasflow path within the lowest pressure compressor 14. The further higherpressure stages may be spaced axially along the gas flow path throughthe compressor in a downstream (rearward) direction.

The lowest pressure turbine 19 similarly comprises one or more turbinestages. In the embodiment shown in FIG. 1, the lowest pressure turbine19 comprises one stage. Each turbine stage may comprise a row of rotorblades 19 b and a row of stator vanes 19 a, 19 c, which may be variablestator vanes (in that their angle of incidence may be variable). The rowof rotor blades and the row of stator vanes may be axially offset fromeach other.

The one or more turbine stages forming the lowest pressure turbine 19may comprise a highest pressure stage, and may further comprise one ormore turbine stages of decreasing pressure to a lowest turbine pressurestage. The lowest pressure turbine stage may be located furthestdownstream within the lowest pressure turbine 19. The further pressurestages are spaced axially in an upstream (forward) direction along thegas flow path through the turbine. In embodiments with only one stage,the single stage is the lowest pressure stage.

Each row of rotor blades provided in the lowest pressure compressor 14and the lowest pressure turbine 19 may form an annular array of rotorblades 44 carried by a respective rotor hub 46 (or rotor disc), as shownby way of example in FIGS. 15A and 15B. Each of the rotor blades 44 maybe coupled to the hub 46 via a root received in a corresponding slot ina peripheral edge of the hub. Each rotor blade 44 may be defined ashaving a radial span extending from the root 46 (or hub) at a radiallyinner gas-washed location, or 0% span position, to an outer most radialtip 48 at a 100% span position. The radius at the hub and the radius atthe tip may both be measured at the leading edge (or axiallyforwardmost) part of the rotor blade. The radial span each rotor blade44 refers to the gas-washed portion of the rotor blade, i.e. the portionradially outside any platform at which it is coupled to the hub.

Each of the rotor blades 44 forming the compressor or turbine stages 19b may have a leading edge mean radius point (or mid blade span) and atrailing edge mean radius point. The mean radius point is defined as themidpoint between the 0% span position and the 100% span position. It maybe measured at the rotor blade leading edge (axially forward-most edge)or trailing edge (axially rearward-most edge) to give the leading edgemean radius point and the trailing edge mean radius point respectively.

The fan 23 comprises an annular array of fan blades 64 extending from ahub 66. Each fan blade 64 may be defined as having a radial spanextending from a root 66 received in a slot in the fan hub 66 at aradially inner gas-washed location, or 0% span position, to a tip 68 ata 100% span position. The ratio of the radius of the fan blade 64 at thehub to the radius of the fan blade at the tip may be less than (or onthe order of) any of: 0.4, 0.39, 0.38 0.37, 0.36, 0.35, 0.34, 0.33,0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. The ratio of theradius of the fan blade 64 at the fan hub 66 to the radius of the fanblade at the tip 68 may be in an inclusive range bounded by any two ofthe values in the previous sentence (i.e. the values may form upper orlower bounds). These ratios may commonly be referred to as thehub-to-tip ratio. The fan blade 64 has a leading edge 64 a and atrailing edge 64 b defined along the direction of gas flow through theengine. The radius at the fan hub 66 and the radius at the tip 68 mayboth be measured at the leading edge 64 a (or axially forward-most) partof the blade. The hub-to-tip ratio refers to the gas-washed portion ofthe fan blade, i.e. the portion radially outside any platform by whicheach fan blade is coupled to the hub.

The gas turbine engine may be described by one or more of the followingparameters:

Engine Length:

Referring to FIGS. 5A and 5B, the gas turbine engine 10 of theembodiments being described has an engine length (labelled 110 in thefigures) defined as the axial distance between: the intersection of theleading edge 64 a of one of the fan blades 64 and the hub 66; and thetrailing edge mean radius point of one of the rotor blades 44 providedin the lowest pressure stage 19 b of the lowest pressor turbine 19.

In the embodiments being described, the engine length 110 is in therange from 200 cm to 500 cm, and more particularly from 300 cm to 450cm. In an embodiment comprising a fan 23 with a fan tip radius 102 inthe range from 110 cm to 150 cm, the engine length 110 may be in therange from 300 cm to 360 cm. In an embodiment comprising a fan 23 with afan tip radius 102 in the range from 155 cm to 200 cm, the engine length110 may be in the range from 370 cm to 470 cm, or from 390 cm to 470 cm.

Core Length:

Referring to FIGS. 4A and 4B, the gas turbine engine 10 has a corelength 104 defined as the axial distance between a forward region of thelow pressure compressor 14 and a rearward region of the low pressureturbine 19, and more specifically the axial distance between the meanradius point (mid blade span) of the first stage of the low pressurecompressor 14 blade leading edge and the mean radius point (mid bladespan) of the lowest pressure turbine rotor stage 19 b blade trailingedge of the low pressure turbine 19.

The first stage of the low pressure compressor 14 is shown in black inFIG. 4A, at the forward end of the core length 104. The lowest pressureturbine rotor stage 19 b of the low pressure turbine 19 is also shown inblack, at the rearward end of the core length 104.

In the embodiments being described, the core length 104 is measuredalong a centreline 9 of the engine 10 from a mean radius point of thefirst stage of the compressor blade leading edge to a mean radius pointof the lowest pressure turbine rotor stage 19 b blade trailing edge ofthe turbine 19.

The core length is in the range from 150 cm to 350 cm in the embodimentbeing described, and more specifically in the range from 160 cm to 320cm. In an embodiment comprising a fan 23 with a fan tip radius in therange from 110 cm to 150 cm, the core length may be in the range from160 cm to 260 cm. In an embodiment comprising a fan 23 with a fan tipradius in the range from 155 cm to 200 cm, the core length may be in therange from 240 cm to 320 cm.

Fan Tip Radius:

The radius 102 of the fan 23, also referred to as the fan tip radius102, or R_(fan tip), may be measured between the engine centreline 9 andthe tip 68 a of a fan blade 64 at its leading edge 64 a (in a radialdirection). The fan diameter may simply be defined as twice the radius102 of the fan 23.

In the embodiments being described, the fan tip radius 102 is in therange from 95 cm to 200 cm, or from 110 cm to 200 cm. In someembodiments, the fan tip radius is in the range from 95 cm to 150 cm orfrom 110 cm to 150 cm. In some alternative embodiments, the fan tipradius is in the range from 155 cm to 200 cm

In some embodiments, the fan diameter is in the range from 190 cm to 300cm, or 220 cm to 300 cm. In some alternative embodiments, the fandiameter is in the range from 310 cm to 400 cm.

Fan Face Area:

The fan face area, A_(fan face), is defined as the circular area sweptby the fan blade tips 68 at the axial position of the fan blade leadingedge 64 a tip. The fan face area is measured in a radial plane. Theskilled person will appreciate that A_(fan face) is at leastsubstantially equivalent to the area within the inner surface of thenacelle 21 at the axial position of the leading edge blade tips (as theblade tip leading edges are arranged to lie very close to the innersurface of the nacelle) for the engine 10 being described.

In the embodiment being described, nacelle inner radius at the axialposition of the leading edge blade tips 68 a is arranged to be slightlylarger than the fan tip radius 102, such that the fan 23 can fit withinthe nacelle 21 without the blade tips 68 rubbing the nacelle 21. Moreparticularly, in the embodiment being described the engine 10 comprisesan engine fancasing 21 a adjacent the blade tips 68 a; the nacelle 21 ismounted on/around the engine fancasing 21 a such that the enginefancasing 21 a effectively forms an inner part of the nacelle 21 onceassembled. The fan casing inner radius at the axial position of theleading edge blade tips 68 a is arranged to be slightly larger than thefan tip radius 102, such that the fan 23 can fit within the enginefancasing 21 a without the blade tips 68 rubbing the fan casing 21 a. Inthe embodiments shown in the Figures, the engine fancasing 21 a extendsonly in the region of the fan 23. In alternative embodiments, thefancasing 21 a may extend rearwardly, for example to the axial locationof a bypass duct outlet guide vane (OGV) 58.

In use, fan geometry may change, for example due to aerodynamic andcentrifugal running loads—the fan 23 may expand more than the nacelle 21and/or more than the fancasing 21 a; the nacelle inner radius maytherefore be selected to accommodate the fan 23 in its expanded state.The skilled person would appreciate that the change in fan radius 102 isrelatively small compared to the total fan radius, for example beingaround 0.1-3 mm for a radius of 95 cm or above, and that values for theratios disclosed herein are therefore not substantially affected bywhether fan radius 102 is measured when cold, or taken in use, or indeedwhether nacelle inner radius at the axial position of the fan blade tipleading edges is used in the place of a measurement of the radius of thefan 23 itself.

The fan face area may be defined as follows:A _(fan face) =πR _(fan tip) ²Where R_(fan tip) is the radius 102 of the fan 23 at the leading edge(i.e. at the tips 68 a of the leading edge 64 a of the fan blades 64).

In the embodiment being described, the area is defined in a radial plane(at the axial location of the leading edge tip 68 a), and can thereforebe calculated using the fan tip radius 102. In alternative embodiments,fan blade curvature may be taken into account when calculating fan facearea.

In some embodiments, the fan diameter is in the range from 220 cm to 300cm and the fan face area is in the range of 2.8 m² to 7.1 m². In somealternative embodiments, the fan diameter is in the range from 310 cm to400 cm and the fan face area is in the range of 7.5 m² to 12.6 m².

Fan Flow Area:

The fan flow area, A_(flow), is defined as the annular area between fanblade tips 68 and the hub 66 at the axial position of the fan bladeleading edge tip 68 a. The fan flow area is measured in a radial plane.The skilled person will appreciate that A_(flow) is at leastsubstantially equivalent to the area of the annulus formed between thehub 66 of the fan 23 and the inner surface of the nacelle 21 immediatelyadjacent the leading edge blade tips (as the blade tip leading edges 64a are arranged to lie very close to the inner surface of the nacelle21—noting the above comments about the fancasing 21 a) for the fanengine 10 being described, and is therefore equivalent to the fan facearea minus the area taken by the hub 66.

As referred to herein, the flow area of the fan (A_(flow)) is definedas:A _(flow)=π(R _(fan tip) ² −R _(hub) ²)Where:R_(fan tip) is the radius 102 (in metres) of the fan 23 at the leadingedge (i.e. at the tips 68 a of the leading edge of the fan blades 64);R_(hub) is the distance 103 (in metres) between the centreline of theengine and the radially inner point on the leading edge of the fan blade(i.e. of radially inner point of the gas-washed surface of the fanblade)—this is equivalent to the radius of the hub 66 of the fan 23 atthe point at which the leading edge of each blade 64 is connectedthereto, and may be referred to as the hub radius.

In one embodiment, the ratio of the radius of fan blade 64 at its hub 66to the radius of the fan blade at its tip 68 may be less than 0.33.

In the embodiment being described, the flow area is defined in a radialplane, and can therefore be calculated using the fan tip radius 102 andthe hub radius 103.

Position of Centre of Gravity:

The gas turbine engine 10 has a position of centre of gravity (CoG)(labelled as 108 in FIGS. 5A and 5B) defined as the axial distancebetween: the intersection of a leading edge 64 a of one of the fanblades 64 and the fan hub 66; and the centre of gravity of the engine10. The centre of gravity may be measured for the engine 10 includingthe nacelle 21 and any components it surrounds, and does not include anyattaching hardware (such as a pylon 53) provided to mount the nacelle 21or other support structure.

In the embodiment being described, the CoG position is between 100 cmand 230 cm from the intersection of a leading edge 64 a of one of thefan blades 64 and the fan hub 66.

In some embodiments, the fan diameter is in the range from 220 cm to 300cm (i.e. a fan tip radius in the range from 110 cm to 150 cm) and theCoG position is in the range from 140 cm to 180 cm. In some alternativeembodiments, the fan diameter is in the range from 310 cm to 400 cm(i.e. fan tip radius of 155 cm to 200 cm) and the CoG position is in therange from 160 cm to 230 cm.

Gearbox Location:

In embodiments with a gearbox 30, the gas turbine 10 has a gearboxlocation (labelled as 112 in FIGS. 6A and 6B) corresponding to arelative position of the gearbox 30 along the engine length 110. Thegearbox location 112 may be measured between: the intersection of aleading edge 64 a of one of the fan blades 64 and the hub 66; and aradial centre plane of the gearbox 30, the radial centre plane being atthe midpoint between the front face of a most forward gear mesh of thegearbox and the rear face of a most rearward gear mesh of the gearbox.In the embodiment being described having an epicyclic gearbox 30, thegearbox location 112 may be defined as the axial distance between: theintersection of a leading edge 64 a of one of the fan blades 64 and thehub 66; and a radial plane intersecting the axial centre point of thering gear 38 of the gearbox 30.

In the embodiment being described, the gearbox location is between 50 cmand 110 cm from the intersection of a leading edge 64 a of one of thefan blades 64 and the fan hub 66.

In some embodiments, the fan diameter is in the range from 220 cm to 300cm and the gearbox location 112 is in the range from 50 cm to 80 cm. Insome alternative embodiments, the fan diameter is in the range from 310cm to 400 cm and the gearbox position 112 is in the range from 80 cm to110 cm.

Diameter of Turbine at Lowest Pressure Rotor Stage (Turbine Diameter):

Referring to FIGS. 11A and 11B, the gas turbine engine 10 has a diameter122 of the low pressure turbine 19 at its lowest pressure rotor stage 19b. This may be referred to as the “turbine diameter” herein. The skilledperson will appreciate that the diameter of the turbine 19 may varyalong the length of the turbine 19, and that a particular axial position(in this case that of the lowest pressure rotor stage 19 b) is thereforeidentified to define a specific diameter value.

The skilled person would appreciate that the lowest pressure rotor stage19 b is the rearmost rotor stage of the turbine 19, and that therearmost rotor stage 19 b of the turbine 19 world be referred to as thelowest pressure rotor stage of the turbine 19 even when the engine 10 isnot in use; i.e. even when pressure does not vary substantially acrossthe engine.

In the embodiment being described the diameter 122 at the lowestpressure rotor stage 19 b is measured at the axial location of blade tiptrailing edges of rotor blades 44 of the lowest pressure rotor stage 19b. The turbine diameter 122 is defined as the diameter at the point ofintersection between the lowest pressure rotor stage blade 44 trailingedge and the outer edge of the gas path annulus.

On a shrouded turbine blade 44 such as that of the embodiment beingdescribed (illustrated in FIG. 15B), the underside of the shroud 49defines the turbine diameter 122 (where “underside” is defined as thesurface of the shroud closest to engine centre), as the shroud 49provides an edge to the gas path annulus Whilst the blade 44 extendsinto the shroud 49 in the embodiment being described, so as tofacilitate mounting of the shroud 49 on the blades 44, the point atwhich the blade 44 enters the shroud 49 may be thought of as the bladetip 48, as it is the radially outermost part of the blade 44 exposed tothe gas flow. On a shroudless rotor 19 b′, i.e. a turbine 19 without ashroud mounted on the blades, such as that illustrated in FIG. 15A, thetips 48 of the blades 44 define the diameter 122′.

In the embodiment shown in FIGS. 4A, 11A and 11B, the turbine 19 hasonly one rotor 19 b (i.e. one row of rotor blades 44, at a specificaxial location), and so the only rotor of the turbine 19 is the lowestpressure rotor stage of the turbine. The rotor 19 b is located betweentwo stators 19 a, 19 c. The rearmost stator 19 c may also be referred toas an Outlet Guide Vane (OGV). In alternative embodiments, multiplerotors may be present within the turbine 19. The lowest pressure rotorstage of the turbine 19 b is the rearmost rotor stage of the turbine 19in such embodiments, as the skilled person will appreciate that pressuredecreases along the length of the turbine 19, from front to back. Insuch embodiments, the turbine 19 comprises a plurality of rotor stagesincluding a lowest pressure rotor stage located furthest downstream.

The turbine 19 of the embodiment being described comprises a turbinecasing. The rotor(s) 19 b and stators 19 a, 19 c are mounted within thecasing. In the embodiment being described the turbine diameter 122 is atleast substantially equal to an inner diameter of the turbinecasing—i.e. shroud width and/or blade tip to casing spacing is smallrelative to the turbine diameter 122. In the embodiment being described,the engine 10 comprises a casing 11 a around the engine core 11, and theturbine casing is provided by a part of the core casing 11 a. Inalternative embodiments, the turbine casing may be separate.

The diameter 122 of the low pressure turbine 19 at its lowest pressurerotor stage 19 b (shown in FIG. 11A) is equal to twice the radius 106 ofthe low pressure turbine 19 at its lowest pressure rotor stage (shown inFIG. 4A). The radius 106 of the low pressure turbine 19 at its lowestpressure rotor stage 19 b is the distance between the engine centrelineand the point of intersection between the lowest pressure rotor stageblade trailing edge and the outer edge of the gas path annulus (which isthe underside of the shroud for the shrouded rotor of the embodimentbeing described, but would be defined by the blade tips in embodimentswith an unshrouded rotor).

In the embodiment being described, the turbine diameter 122 at thelowest pressure rotor stage is in the range from 70 cm to 170 cm. Inembodiments with an engine 10 with a fan tip radius 102 in the rangefrom 110 cm to 150 cm, the turbine diameter 122 at the lowest pressurerotor stage may be in the range from 100 cm to 120 cm. In embodimentswith an engine 10 with a fan tip radius 102 in the range from 155 cm to200 cm, the turbine diameter 122 at the lowest pressure rotor stage maybe in the range from 120 cm to 170 cm.

Bypass Exhaust Nozzle Outer Radius:

The bypass duct 22 has a bypass exhaust nozzle 18—as air drawn inthrough the fan 23 and bypassing the core 11 passes through the bypassduct 22 and out of the bypass exhaust nozzle 18, the bypass exhaustnozzle may be referred to as a fan nozzle 18. The bypass exhaust nozzleouter radius 114 may therefore be referred to as the fan nozzle outerradius 114. In the embodiment being described, an inner surface of thenacelle 21 defines the outer surface of the bypass exhaust nozzle 18.

The bypass exhaust nozzle outer radius 114, shown in FIGS. 7A to C, isdefined as the radius at the outer edge of the bypass exhaust nozzleexit. The radius 114 is measured from the engine centre line 9 to therearmost tip 21 b of the inner surface of the nacelle 21, in a radialplane. The radial plane may be referred to as an exit plane 54 of thebypass exhaust nozzle 18. The bypass exhaust nozzle 18 ends where thenacelle 21 ends, making the rearmost tip 21 b of the nacelle 21 theaxial position of the exit from the bypass exhaust nozzle 18 in theembodiment being described.

In the embodiments being described, the outer radius 114 of the bypassexhaust nozzle 18 is in the range of 100 cm to 200 cm, and particularlyin the range 100 cm to 190 cm. In embodiments with an engine 10 with afan tip radius 102 in the range from 110 cm to 150 cm, the outer radius114 of the bypass exhaust nozzle 18 may be in the range from 100 cm to145 cm. In embodiments with an engine 10 with a fan tip radius 102 inthe range from 155 cm to 200 cm, the outer radius 114 of the bypassexhaust nozzle 18 may be in the range from 145 cm to 190 cm.

Bypass Exhaust Nozzle Inner Radius:

The bypass exhaust nozzle inner radius 116 may also be referred to asthe fan nozzle inner radius 116. The bypass exhaust nozzle inner radius116 is defined as the radius at the inner edge of the bypass exhaustnozzle exit. The radius 116 is measured from the engine centre line 9 tothe point on the engine core 11 in the same axial position as therearmost tip 21 b of the inner surface of the nacelle 21, in the sameradial plane (which may be referred to as the exit plane 54 of thebypass exhaust nozzle 18). The bypass exhaust nozzle 18 ends where thenacelle 21 ends, meaning that the rearmost tip of the nacelle 21 definesthe axial position of the exit from the bypass exhaust nozzle 18.

In the embodiments being described, the inner radius 116 of the bypassexhaust nozzle 18 is in the range from 50 cm to 125 cm, and optionallyfrom 65 cm to 110 cm. In embodiments comprising an engine 10 with a fantip radius 102 in the range from 110 cm to 150 cm, the inner radius 116of the bypass exhaust nozzle 18 may be in the range from 65 cm to 90 cm.In embodiments comprising an engine 10 with a fan tip radius 102 in therange from 155 cm to 200 cm, the inner radius 116 of the bypass exhaustnozzle 18 may be in the range from 80 cm to 110 cm.

Bypass Exhaust Nozzle Flow Area:

The flow area A_(b) of the bypass exhaust nozzle 18 at the nozzle exitmay be defined as shown in FIG. 3B. A minimum distance R_(b) across thenozzle 18 experienced by the bypass gas flow (B) is identified bysuperposing a circle C_(b) with a centre point CP_(b) at the rearmosttip of the nacelle 21 and expanding that circle until it makes contactwith the inner annulus line of the bypass duct 22 (i.e. the outersurface of the engine core 11).

The area, A_(b), experienced by the flow is defined based on rotatingthat minimum distance R_(b) around the circumference, so forming anangled, approximately annular area, and subtracting the blocked area (inthis embodiment blocked by the pylon 53, as illustrated on theright-hand side of FIG. 3B, which shows a view (not to scale) of thenozzle area in a rearward-facing radial plane).

The skilled person would appreciate that, in the embodiment shown, theminimum distance R_(b) across the nozzle 18 is angled with respect tothe radius of the engine 10, i.e. not perpendicular to the enginecentreline 9, and that the minimum distance R_(b) is therefore not equalto the difference between the inner radius 116 of the bypass exhaustnozzle 18 and the outer radius 114 of the bypass exhaust nozzle 18, andis not measured in the same plane as those radii. In alternativeembodiments, the circle C_(b) may make contact with the engine core 11at the same axial position as the rearmost tip of the nacelle 21—in suchembodiments, the minimum distance R_(b) would be equal to the differencebetween the inner radius 116 of the bypass exhaust nozzle 18 and theouter radius 114 of the bypass exhaust nozzle 18.

The flow area, A_(b) of the bypass exhaust nozzle at the bypass ductexhaust nozzle exit is between 1.9 m² and 5.8 m² in the embodiment beingdescribed. In embodiments with an engine 10 with a fan tip radius 102 inthe range from 110 cm to 150 cm, the flow area of the bypass ductexhaust nozzle at the bypass duct nozzle exit may be in the range from1.9 m² to 4.5 m². In embodiments with an engine 10 with a fan tip radius102 in the range from 155 cm to 200 cm, the flow area of the bypass ductexhaust nozzle at the bypass duct exhaust nozzle exit may be in therange from 4.5 m² to 5.8 m².

Core Exhaust Nozzle Flow Area:

The flow area A_(c) of the core exhaust nozzle 20 at the nozzle exit maybe defined as for that of the bypass exhaust nozzle 18, as shown in FIG.3B. A minimum distance R_(c) across the nozzle 20 experienced by the gasflow is identified by superposing a circle C_(c) with a centre pointCP_(c) at the rearmost tip of the engine core casing/inner fixedstructure 11 a, and expanding that circle until it makes contact withthe inner annulus line of the core nozzle 20 (i.e. the outer surface ofthe exhaust cone 67).

The area, A_(c), experienced by the flow is then defined based onrotating that minimum distance R_(c) around the circumference, soforming an angled, approximately annular area, and subtracting theblocked area (in this embodiment blocked by the pylon 53, as illustratedon the right-hand side of FIG. 3B, which shows a view of the nozzle areain a rearward-facing radial plane). In the embodiment shown, the minimumdistance R_(c) across the nozzle 20 is angled with respect to the radiusof the engine 10, i.e. not perpendicular to the engine centreline 9. Inalternative embodiments, the minimum distance R_(c) may be a radialdistance.

In the embodiment being described, the flow area, A_(c), of the coreexhaust nozzle at the core exhaust nozzle exit is between 0.4 m² and 1.3m². In embodiments with an engine (10) with a fan tip radius (102) inthe range from 110 cm to 150 cm, the flow area of the core exhaustnozzle at the core exhaust nozzle exit may be in the range from 0.4 m²to 0.6 m². In embodiments with an engine (10) with a fan tip radius(102) in the range from 155 cm to 200 cm, the flow area of the coreexhaust nozzle at the core exhaust nozzle exit may be in the range from0.6 m² to 1.3 m².

Outer Bypass Duct Wall Angle:

The bypass duct 22 is partly defined by an outer wall formed by theinner surface of the nacelle 21 as illustrated in FIG. 12A. In thisembodiment, a bypass duct outlet guide vane (OGV) 58 is provided thatextends radially across the bypass duct 22 between an outer surface ofthe engine core 11 (e.g. the core casing 11 a) and the inner surface ofthe nacelle 21. The OGV extends between a radially inner tip 58 a and aradially outer tip 58 b (see FIG. 12C) and has a leading (or upstream)edge and a trailing (or downstream) edge relative to the direction ofgas flow B through the bypass duct 22.

An outer wall axis 60 is defined joining the radially outer tip 58 b ofthe trailing edge of the bypass duct outlet guide vane 58 and therearmost tip 21 b of the inner surface of the nacelle 21. The outer wallaxis 60 lies in a longitudinal plane containing the centreline 9 of thegas turbine engine. In the described embodiment, the outer wall axis 60is defined based on fixed nacelle (e.g. fan duct) geometry. The rearmosttip 21 b of the inner surface of the nacelle 21 therefore remains in aconstant position relative to the OGV. In other embodiments, the gasturbine engine 10 may have a variable area fan nozzle as describedabove. In such embodiments, the rearmost tip of the inner surface of thenacelle 21 (and so the outer wall axis 60) may be movable during use ofthe engine. The outer wall axis 60 may be defined based on the positionof the rearmost tip 21 b of the inner surface of the nacelle duringcruise conditions. The cruise conditions may be as described elsewhereherein.

The outer bypass duct (BPD) wall angle 126 is defined by the anglebetween the outer wall axis 60 and the centreline 9 of the engine asillustrated in FIGS. 12A, 12B and 12C. A positive value of the BPD wallangle 126 corresponds to the outer wall axis 60 sloping away from theengine centre line 9 when moving in a rearward direction along the axis,i.e. the rearmost tip of the inner surface of the nacelle 21 is furtherfrom the engine centre line 9 than the radially outer tip of thetrailing edge of the bypass OGV. A positive BPD wall angle isillustrated in FIGS. 12A and 12B. A negative value of the BPD wall anglecorresponds to the outer wall axis 60 sloping towards the enginecentreline 9 when moving rearward along the axis. A negative BPD wallangle is illustrated in FIG. 12C. In this case the rearmost tip of theinner surface of the nacelle 21 is closer to the engine centre line 9than the radially outer tip of the trailing edge of the bypass OGV.

A bypass duct outlet guide vane radius, measured radially between theengine centreline 9 and the radially outer tip 58 b of the trailing edgeof the bypass OGV, may be in a range from 90 cm to 210 cm. For example,for an engine 10 with a fan tip radius 102 in the range from 110 cm to150 cm, the bypass duct outlet guide vane radius may be in the rangefrom 90 cm to 150 cm, or more specifically from 110 cm to 135 cm. For anengine 10 with a fan tip radius 102 in the range from 155 cm to 200 cm,the bypass duct outlet guide vane radius may be in the range from 160 cmto 210 cm, or more specifically from 170 cm to 200 cm.

Fan Axis Angle:

Referring to FIGS. 9A and 9B, the gas turbine engine 10 has a fan axisangle 118 related to the angle between the outer radial tip 68 of thefan blades 64 and the outer radial tip 48 of the rotor blades 44 of thelowest pressure stage 19 b of the low pressure turbine 19. A fan tipaxis 62 lies in a common plane with the engine centreline 9. The fan tipaxis 62 joins the radially outer tip 68 a of the leading edge 64 a ofthe fan blade 64 and the radially outer tip of the trailing edge of oneof the rotor blades 44 of the lowest pressure stage 19 b of the lowpressure turbine 19.

The fan axis angle 118 is defined as the angle between the fan tip axis62 and the engine centreline 9 as illustrated in FIG. 9B.

As described elsewhere herein, the rotor blades 44 of the lowestpressure stage 19 b of the lowest pressure turbine 19 may be shrouded orunshrouded. If the rotor blades 44 are shrouded, the outer radial tip ofthe rotor blades is taken to be the underside of the shroud 49 (whichprovides the edge of the gas-flow annulus). If the rotor blades 44′ areunshrouded, it is the blade tips 48′ of the rotor 19 b′.

Fan-Turbine Radius Difference:

Referring to FIGS. 10A and 10B, the gas turbine engine 10 has afan-turbine radius difference 120 defined as a radial distance between:a point of a circle intersecting (e.g. swept by) the radially outer tip48 of the trailing edge of the rotor blades 44 of the lowest pressurestage 19 b of the low pressure turbine 19; and a point on a circleintersecting (e.g. swept by) the radially outer tip 68 a of the leadingedge 64 a of the fan blades 64.

The fan-turbine radius difference 120 may be in a range from 50 cm to120 cm. The fan-turbine radius difference 120 may be in a range between55 cm to 85 cm, for example for an engine 10 with a fan tip radius 102in the range from 110 cm to 150 cm. The fan-turbine radius difference120 may be in a range between 90 cm to 120 cm, for example for an engine10 with a fan tip radius 102 in the range from 155 cm to 200 cm.

As described above, the rotor blades 44 of the lowest pressure stage 19b of the lowest pressure turbine 19 may be shrouded or unshrouded. Ifthe rotor blades 44 are shrouded, the outer radial tip of the rotorblades 48 is taken to be the underside of the shroud 49 (the edge of thegas-flow annulus). If the rotor blades 44′ are unshrouded, it is theblade tips 48′ of the rotor 19 b′.

Distance Between Ground Plane and Wing:

In the embodiments being described, the distance 124 is measured withrespect to the ground 50, as shown in FIGS. 11A and 11B—a ground planeis defined as the plane 50 on which the aircraft 70 would rest afterlanding/before take-off—e.g. the surface of a runway or floor of ahangar. The skilled person would appreciate that, in most embodiments,aircraft landing gear would be extended and in contact with the groundplane 50. The vertical distance between the ground 50 and the wing 52 ismeasured.

As the wing 52 varies in height along the axial direction in theembodiments being described (due to its aerofoil shape), an axialposition is selected for this measurement—in the embodiment beingdescribed, the axial location of the leading edge 52 a of the wing 52 isselected. In particular, the ground to wing distance 124, as definedherein, is the vertical distance between the ground plane 50 and thecentre point of the wing's leading edge 52 a. The distance 124 betweenthe ground plane 50 and the wing 52 is therefore measured to the centrepoint 52 a of the leading edge 52 a of the wing 52.

As the wing 52 varies in height along its length in the embodiment beingdescribed, from aircraft 70 to wing tip, a location along the length ofthe wing 52 is also selected for the measurement. In the embodimentbeing described, the selected location is directly above the enginecentreline 9 (the engine axis 9). The distance 124 between the groundplane 50 and the wing 52 is therefore measured along a lineperpendicular to the ground plane 50 and passing through, and in atleast this embodiment perpendicular to, an axial centre line of theengine 10.

The skilled person will appreciate that the ground to wing distance 124may also vary depending on loading of the aircraft 70. As used herein,maximum take-off weight (MTOW) is assumed for the definition of theground to wing distance 124.

Maximum Take-Off Weight:

The MTOW of an aircraft 70 may also be referred to as maximum grosstake-off weight (MGTOW) or maximum take-off mass (MTOM) of an aircraft70. The MTOW is the maximum weight at which a pilot is allowed toattempt to take off, due to structural or other limitations. The skilledperson will appreciate that maximum Take-Off Weight (MTOW) for anaircraft 70 is a standard parameter issued with an aircraft'scertification, and that the MTOW can therefore be trivially identifiedfor any commercial aircraft 70, and can be determined according tostandard practices for any aircraft 70.

Downstream Blockage:

Downstream blockage provides a measure of how much of the space beneatha wing 52 of an aircraft 70 is taken up by the gas turbine engine 10. Inthe embodiments being described, the downstream blockage is measuredwith respect to the ground plane 50. Herein, a downstream blockage ratiois defined as:

$\frac{\begin{matrix}{{the}\mspace{14mu}{turbine}\mspace{14mu}{diameter}\mspace{14mu}(122)\mspace{14mu}{at}\mspace{14mu}{an}\mspace{14mu}{axial}\mspace{14mu}{location}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{lowest}} \\{{pressure}\mspace{14mu}{rotor}\mspace{14mu}{stage}\mspace{14mu}\left( {19\mspace{14mu} b} \right)}\end{matrix}}{{distance}\mspace{14mu}{between}\mspace{14mu}{ground}\mspace{14mu}{plane}{\mspace{11mu}\;}{and}\mspace{14mu}{wing}\mspace{14mu}(124)}$The turbine diameter 122 and distance 124 between the ground plane 50and wing 52 are as defined above.Quasi Non-Dimensional Mass Flow Rate (Q):

A quasi non-dimensional mass flow rate Q is defined as:

$Q = {W{\frac{\sqrt{T0}}{P{0.A_{flow}}}.}}$Where:W is mass flow rate through the fan in Kg/s;T0 is average stagnation temperature of the air at the fan face inKelvin;P0 is average stagnation pressure of the air at the fan face in Pa; andA_(flow) is the flow area of the fan in m², as defined above.The parameters W, T0, P0 and A_(flow) are all shown schematically inFIG. 14A.

At cruise conditions of the gas turbine engine 10 (which may be asdefined elsewhere herein), the value of Q is, for example, in the rangeof from 0.029 to 0.036 Kgs⁻¹N⁻¹K^(1/2). In particular, as used herein,cruise conditions may mean cruise conditions of an aircraft 70 to whichthe gas turbine engine 10 is attached. Such cruise conditions may beconventionally defined as the conditions at mid-cruise, for example theconditions experienced by the aircraft 70 and/or engine 10 at themidpoint (in terms of time and/or distance) between top of climb andstart of decent. Purely by way of example, the cruise conditions maycorrespond to: a forward Mach number of 0.8; a pressure of 23000 Pa; anda temperature of −55 deg C.

Also at cruise conditions, the gas turbine engine 10 generates a thrustT (which may be referred to as a cruise thrust), shown schematically inFIG. 14A. This thrust may be equal to the thrust required to maintainthe cruise forward speed of an aircraft 70 to which the gas turbineengine 10 is attached, divided by the number of engines 10 provided tothe aircraft.

At cruise conditions, the thrust, T, divided by the mass flow rate, W,through the engine (which is equal to the mass flow rate W at the faninlet) is, for example, in the range of from 70 Nkg⁻¹s to 110 Nkg⁻¹s.

Pressure Ratio of Bypass Exhaust Nozzle:

The bypass exhaust nozzle 18 may also be referred to as the fan nozzle18. The skilled person will appreciate that a nozzle pressure ratio(NPR) is generally defined as:

$\frac{{total}\mspace{14mu}{pressure}\mspace{14mu}{at}\mspace{14mu}{nozzle}\mspace{14mu}{exit}}{{ambient}\mspace{14mu}{pressure}\mspace{14mu}{of}\mspace{14mu}{surroundings}}$The pressure ratio of the bypass exhaust nozzle 18 is therefore:

$\frac{{total}\mspace{14mu}{pressure}\mspace{14mu}{at}\mspace{14mu}{bypass}\mspace{14mu}{nozzle}\mspace{14mu}{exit}}{{{ambient}\mspace{14mu}{pressure}}\;}$

The location of the exit from the bypass exhaust nozzle 18 is asdescribed above, and as shown in FIGS. 13A and 13B. In particular, anexit plane 54 is defined at the exit of the bypass exhaust nozzle 18.The exit plane 54 is defined as the annular, radial plane extendingacross the bypass exhaust nozzle 18 at the axial location of therearward tip of the nacelle 21. The total pressure at the bypass nozzleexit, P_(BE), is defined at this plane—i.e. the sum of the static anddynamic pressures at the nozzle exit 54 of the bypass exhaust nozzle 18is determined as the total pressure P_(BE).

The skilled person would appreciate that pressures throughout the engine10 may be modelled from aerodynamic principles, and/or one or morepressure sensors (e.g. in the form of a pressure rake) may be locatedwithin the bypass nozzle 18 or elsewhere in the bypass duct 22 to recordactual local pressures, and the pressure at the bypass nozzle exit plane54 may be determined from those measurements.

A known value based on aircraft altitude may be used for the ambientpressure, P_(Amb).

The total pressure used to calculate the bypass exhaust nozzle pressureratio is the total pressure under cruise conditions, as defined above.In particular, as used herein, cruise conditions may mean cruiseconditions of an aircraft 70 to which the gas turbine engine 10 isattached. Such cruise conditions may be conventionally defined as theconditions at mid-cruise, for example the conditions experienced by theaircraft 70 and/or engine 10 at the midpoint (in terms of time and/ordistance) between top of climb and start of decent. Purely by way ofexample, the cruise conditions may correspond to: a forward Mach numberof 0.8; a pressure of 23000 Pa; and a temperature of −55 deg C.

Pressure Ratio of Core Nozzle:

The skilled person will appreciate that a nozzle pressure ratio (NPR) isdefined as:

$\frac{{total}\mspace{14mu}{pressure}\mspace{14mu}{at}\mspace{14mu}{nozzle}\mspace{14mu}{exit}}{{ambient}\mspace{14mu}{pressure}\mspace{14mu}{of}\mspace{14mu}{surroundings}}$The pressure ratio of the core nozzle 20 is therefore:

$\frac{{total}\mspace{14mu}{pressure}\mspace{14mu}{at}\mspace{14mu}{core}\mspace{14mu}{nozzle}\mspace{14mu}{exit}}{{{ambient}\mspace{14mu}{pressure}}\;}$

The location of the exit from the core nozzle 20 is at the axialposition defined by the rearmost tip of the core casing 11 a/inner fixedstructure 11 a, as shown in FIG. 13A. The location of the exit from thecore exhaust nozzle 20 is as described above, and as shown in FIGS. 13Aand 13B. In particular, an exit plane 56 is defined at the exit of thecore exhaust nozzle 20. The exit plane 56 is defined as the annular,radial plane extending across the core exhaust nozzle 20 at the axiallocation of the rearward tip of the core casing 11 a. The total pressureat the core nozzle exit, P_(CE), is defined at this plane—i.e. the sumof the static and dynamic pressures at the nozzle exit 56 of the coreexhaust nozzle 20 is determined as the total pressure P_(CE).

In the embodiment being described, the exit plane 56 for the core nozzle20 is rearward of the exit plane 54 of the bypass exhaust nozzle 18, asthe core casing 11 a extends further rearward than the nacelle 21. Inalternative embodiments, the exit planes 56, 54 may be closer, may becoplanar, or the order of the planes may be reversed.

The skilled person would appreciate that pressures throughout the engine10 may be modelled from aerodynamic principles, and/or one or morepressure sensors (e.g. in the form of a pressure rake) may be locatedwithin the core nozzle 20 to record actual local pressures, and thepressure at the core nozzle exit plane 56 may be determined from thosemeasurements.

A known value based on aircraft altitude may be used for the ambientpressure, P_(Amb). The same value may be used as for the pressure ratioof the bypass exhaust nozzle 18. The skilled person would appreciatethat the same value for ambient pressure is generally used for bothratios.

As referred to herein, the pressure at a plane (for example totalpressure at bypass nozzle exit or total pressure at core nozzle exit)may be taken as the mean value over that plane.

The total pressure used to calculate the core exhaust nozzle pressureratio is the total pressure under cruise conditions, as defined above.In particular, as used herein, cruise conditions may mean cruiseconditions of an aircraft 70 to which the gas turbine engine 10 isattached. Such cruise conditions may be conventionally defined as theconditions at mid-cruise, for example the conditions experienced by theaircraft 70 and/or engine 10 at the midpoint (in terms of time and/ordistance) between top of climb and start of decent. Purely by way ofexample, the cruise conditions may correspond to: a forward Mach numberof 0.8; a pressure of 23000 Pa; and a temperature of −55 deg C.

Maximum Take-Off Fan Rotational Speed

The rotational speed of the fan 23 may vary during use of the gasturbine engine 10. The fan 23 may have a maximum take-off (MTO)rotational speed (e.g. in rpm) corresponding to the maximum speed atwhich it rotates during take-off of an aircraft 70 to which the gasturbine engine 10 is mounted.

The maximum take-off rotational fan speed may be in a range between 1450rpm to 3020 rpm. For an engine 10 with a fan tip radius 102 in the rangefrom 110 cm to 150 cm, the maximum take-off rotational fan speed may bein a range between 2100 rpm to 3020 rpm or 1970 rpm to 3020 rpm. For anengine 10 with a fan tip radius 102 in the range from 155 cm to 200 cm,the maximum take-off rotational fan speed may be in a range between 1450rpm to 1910 rpm.

It has been found that the parameters defined above may be combined inany one or more of the following ratios in order to provide an improvedgas turbine engine:

Engine Area Ratio

An engine area ratio may be defined as:

$\frac{{the}\mspace{14mu}{fan}\mspace{14mu}{face}\mspace{14mu}{{area}\left( A_{{fan}\mspace{14mu}{face}} \right)}}{{the}\mspace{14mu}{turbine}\mspace{14mu}{diameter}\mspace{14mu}(122) \times {the}\mspace{14mu}{core}\mspace{14mu}{length}\mspace{14mu}(104)}$

The turbine diameter 122 as used in this ratio is the diameter 122 ofthe turbine 19 at the axial position of the lowest pressure rotor stage19 b, as defined above. The skilled person would appreciate that thelowest pressure rotor stage 19 b is the rearmost rotor stage of theturbine 19, and that the rearmost rotor stage 19 b of the turbine 19would be referred to as the lowest pressure rotor stage of the turbine19 even when the engine 10 is not in use; i.e. even when pressure doesnot vary substantially across the engine.

The fan face area (A_(fan face) as defined above) may be thought of asproviding an indication of an area of the engine 10 in a radial plane.The turbine diameter 122 multiplied by the core length 104 may bethought of as an effective area of the engine core 11 in an axial plane.

The skilled person would appreciate that having a larger fan tip radius102, and therefore a larger fan face area, may improve propulsiveefficiency, for example for a given thrust level. An increase of fanradius is illustrated by arrows 23 a in FIG. 4C. Such an increase wouldincrease the engine area ratio if the engine core 11 were unchanged, orhave no effect on the engine area ratio were the core 11 scaled to matchthe larger fan 23. However, the skilled person would appreciate that anengine 10 simply scaled up for a larger fan 23 could potentiallyincrease drag and difficulty of installation, for example increasingdownstream blockage.

In the embodiments being described, the engine core 11 is made smallerthan it would be if simply scaled up for a larger fan 23, so reducingthe engine area ratio. The skilled person would appreciate that reducingthe core size may comprise reducing the core length 104, as illustratedby arrow 11A in FIG. 4C, reducing the turbine diameter 122, asillustrated by arrow 11B in FIG. 4C, or reducing both, as illustrated byarrow 11C in FIG. 4C. The skilled person will appreciate that corelength 104 and diameter 122 may be traded off against each other tooptimally reduce engine core size given various constraints.

In the embodiment being described, both core length 104 and turbinediameter 122 are reduced relative to fan radius 102 as the size of thefan 23 is increased, as compared to known engines. The engine area ratiois therefore higher than that of current aircraft engines.

In the embodiment being described, the engine area ratio is in the rangefrom 1.7 to 3, more particularly in the range from 1.70 to 3.00. In theembodiment being described, the engine area ratio is greater than 1.70.In the embodiment being described, the engine area ratio is in the rangefrom 1.9 to 3, more particularly in the range from 2 to 3, and moreparticularly in the range from 2.1 to 2.5. In various embodiments, thefan tip radius 102 is in the range from 110 cm to 150 cm, and the enginearea ratio is in the range from 1.7 to 2.7. In alternative embodiments,the fan tip radius 102 is in the range from 155 cm to 200 cm, andoptionally wherein the engine area ratio is in the range from 2 to 3. Inthe embodiment being described with respect to FIG. 4A, the fan tipradius 102 is greater than 170 cm.

In the embodiment being described, the turbine diameter 122 varies alongthe length of the turbine 19. In the embodiment being described, theturbine diameter 122 at the lowest pressure rotor stage 19 b has a valuein one or more of the absolute ranges defined above for turbinediameter.

In the embodiment being described, the ratio of the fan tip radius 102to the turbine diameter 122 at the lowest pressure rotor stage 19 b isin the range of 0.8 to 2.1, inclusive.

In the embodiment being described, the engine core length 104 has avalue in one or more of the absolute ranges defined above for corelength.

In the embodiment being described, the ratio of the fan tip radius 102to the core length 104 is in the range of 0.3 to 1.

In the embodiment being described, the gas turbine engine 10 comprises agearbox 30 connected between the core shaft 26 and the fan 23, thegearbox 30 being arranged to receive an input from the core shaft 26 andto provide an output to drive the fan 23 at a lower rotational speedthan the core shaft 26. In alternative embodiments, there may be nogearbox. In the embodiment being described, the gearbox has a gear ratioin the range of from 3 to 5, and more particularly in the range of from3.2 to 3.8.

In the embodiment being described, the turbine 19 is a first turbine 19and the engine 10 comprises a second turbine 17 arranged to rotate at ahigher rotational speed. In alternative embodiments, only one turbine19, or more than two turbines 17, 19 may be present.

In the embodiment being described, a fan axis angle 118 is defined asdescribed above. The fan axis angle 118 is defined as the angle betweenthe fan tip axis 62 and the centreline 9 of the engine as shown in FIGS.9A and 9B. A positive value of the fan axis angle 118 corresponds to thefan tip axis 62 sloping towards the engine centreline 9 when moving in arearward direction along the axis as illustrated in FIG. 9B, i.e. theradially outer tip 68 a of the leading edge 64 a the plurality of fanblades 64 is further from the engine centreline 9 compared to theradially outer tip of the trailing edge of the rotor blades 19 a of thelowest pressure stage of the turbine 19.

In the embodiment being described, the fan axis angle is in a rangebetween 10 to 20 degrees. By providing a fan axis angle 118 in thisrange the gas turbine engine 10 may have a large fan diameter to provideimproved propulsive efficiency, whilst also having a relatively smalldiameter core 11. In the embodiment being described, the fan axis angle118 is in a range from 12 degrees to 16 degrees, more specifically from13 or 14 to 15 degrees, and particularly around 14.5 degrees.

Bypass to Core Ratio

A bypass to core ratio may be defined as:

$\frac{{bypass}\mspace{14mu}{exhaust}\mspace{14mu}{nozzle}\mspace{14mu}{pressure}\mspace{14mu}{ratio}}{{core}\mspace{14mu}{exhaust}\mspace{14mu}{nozzle}\mspace{14mu}{pressure}\mspace{14mu}{ratio}}$

In the embodiment being described with respect to FIGS. 13A and 13B, thebypass to core ratio is configured to be in the range from 1.1 to 2under aircraft cruise conditions, and more specifically in the rangefrom 1.1 to 2.0 and more specifically from 1.10 to 2.00.

In alternative or additional embodiments, the bypass to core ratio mayfall within one or more of the following ranges under aircraft cruiseconditions: from 1.10 to 2.00; above 1.15; and/or from 1.2 to 1.5. Inembodiments with an engine 10 with a fan tip radius 102 in the rangefrom 110 cm to 150 cm, the bypass to core ratio may be in the range from1.0 to 1.4 or from 1.1 to 1.3. In embodiments with an engine 10 with afan tip radius 102 in the range from 155 cm to 200 cm, the bypass tocore ratio may be in the range from 1.3 to 1.6.

In the embodiment being described, the bypass to core ratio can besimplified to the below, as described above, which may be referred to asan extraction ration:

$\frac{{total}\mspace{14mu}{pressure}\mspace{14mu}{at}\mspace{14mu}{bypass}\mspace{14mu}{exhaust}\mspace{14mu}{nozzle}\mspace{14mu}{exit}\mspace{14mu}(54)}{{total}\mspace{14mu}{pressure}\mspace{14mu}{at}\mspace{14mu}{core}\mspace{14mu}{exhaust}\mspace{14mu}{nozzle}\mspace{14mu}{exit}\mspace{14mu}(56)}$

The total pressures at the bypass nozzle exit 54 and core nozzle exit 56may be defined and determined as described above.

In the embodiment being described, the engine core 11 comprises a casing11 a located radially between the core nozzle 20 and the bypass duct 22.In the embodiment being described, an outer surface of the casing 11 aprovides an inner surface of the bypass exhaust duct 22 and bypassnozzle 18, and an inner surface of the casing 11 a provides an outersurface of the core nozzle 20.

In the embodiment being described, the bypass ratio at cruise conditionsis in the range of from 11 to 20, and more particularly in the rangefrom 13 to 20 or 14 to 20.

In the embodiment being described, the core nozzle exit is defined as anexit plane 56 of the core exhaust nozzle 20 (for the purpose of definingpressures), the exit plane 56 extending from a rearmost point of theengine core casing 11 a towards a centreline of the engine 10. In theembodiment being described, the exit plane 56 is defined as a radialplane, perpendicular to an axis of the engine 10, for the purpose ofdefining pressures.

In the embodiment being described, the bypass nozzle exit is defined asan exit plane 54 of the bypass duct exhaust nozzle 18 (for the purposeof defining pressures), the exit plane 54 extending from a rearmostpoint of the nacelle 21 towards a centreline of the engine 10. In theembodiment being described, the exit plane 54 is defined as a radialplane, perpendicular to an axis of the engine 10, for the purpose ofdefining pressures.

In the embodiment being described, the diameter of the bypass exhaustnozzle 18 at the bypass exhaust nozzle exit 54 has a value in one ormore of the absolute ranges defined above for bypass exhaust nozzlediameter.

In the embodiment being described, the flow area A_(b) of the bypassexhaust nozzle 18 at the bypass exhaust nozzle exit 54 is in the rangefrom 2 m² to 6 m², and more particularly from 1.9 m² to 5.8 m². In theembodiment being described, the flow area A_(c) of the core exhaustnozzle 20 at the core exhaust nozzle exit 56 is in the range from 0.4 m²to 1.3 m². In the embodiment being described, the flow areas aremeasured in a plane at an angle to the radial exit planes 54, 56. Inalternative embodiments, the flow areas A_(b), A_(c) may or may not bein the exit planes 54, 56 depending on the angle of the minimum distanceR_(b), R_(c) as discussed above with respect to FIG. 3B.

In the embodiment being described, a ratio of bypass exhaust nozzle 18flow area at the bypass exhaust nozzle exit 54 to flow area of the coreexhaust nozzle 20 at the core exhaust nozzle exit 56 is in the rangefrom 4 to 6, and more particularly in the range from 5 to 6.

In the embodiment being described, the bypass exhaust nozzle 18 and thecore exhaust nozzle 20 are both convergent nozzles. In alternativeembodiments, one or both of the bypass exhaust nozzle 18 and the coreexhaust nozzle 20 may be convergent-divergent nozzles.

In the embodiment being described, the gas turbine engine 10 furthercomprises a gearbox 30 connected between the core shaft 26 and the fan23, the gearbox 30 being arranged to receive an input from the coreshaft 26 and provide an output to drive the fan at a lower rotationalspeed than the core shaft 26. In the embodiment being described, thegearbox 30 has a gear ratio in the range of from 3 to 5, and moreparticularly of 3.2 to 3.8. In alternative embodiments, no gearbox maybe provided or the gear ratio may differ.

In the embodiment being described, the fan tip radius 102 is greaterthan 170 cm. In alternative or additional embodiments, the fan tipradius 102 may be greater than or on the order of any of: 110 cm, 115cm, 120 cm, 125 cm, 130 cm, 135 cm, 140 cm, 145 cm, 150 cm, 155 cm, 160cm, 165 cm, 170 cm, 175 cm, 180 cm, 185 cm, 190 cm or 195 cm.

In the embodiment being described, the fan 23 is particularly large; theskilled person would appreciate that the larger fan 23 may facilitatethe larger pressure difference between the bypass and core exhaustnozzles 18, 20, provided that other engine parameters are adjustedappropriately. In alternative embodiments, the fan 23 may not berelatively large and other engine parameters may be adjusted to providethe desired ratio of pressures.

In the embodiment being described, the turbine 19 is a first turbine 19and the engine 10 comprises a second turbine 17 arranged to rotate at ahigher rotational speed. In alternative or additional embodiments, theengine 10 may only have a single turbine 19, or may have more than twoturbines 17, 19, for example having three or four turbines.

FIG. 13C illustrates a method 1300 of operating an aircraft 70comprising a gas turbine engine 10 as described above. The methodcomprises taking off 1302, reaching cruise conditions 1304, andcontrolling 1306 the aircraft 70 such that the bypass to core ratioremains in the range from 1.1 to 2 during cruise.

The bypass to core ratio may more specifically be within any of theranges defined above. The method 1300 may include controlling the gasturbine engine 10 according to any of the other parameters definedherein.

Engine Length to CoG Ratio

A centre of gravity (CoG) position ratio may be defined as:

-   -   the centre of gravity position (108)/the engine length (110).

The engine length 110 may be measured as the axial distance between aforward region of the fan 23 and a rearward region of the lowestpressure turbine 19. In the embodiment being described, the enginelength 110 is measured as the axial distance between: the intersectionof the leading edge 64 a of one of the plurality of fan blades 64 andthe hub 66; and a mean radius point of the trailing edge of one of therotor blades 44 of the lowest pressure turbine stage of the turbine 19 bas defined above. The mean radius point is the midpoint between a 0%span position and a 100% span position of the rotor blade 44.

In the embodiment being described the gas turbine engine 10 has a singleturbine 19 referred to as the lowest pressure turbine 19. In otherembodiments, a plurality of turbines may be provided. The engine length110 is measured to a rotor of the lowest pressure stage 19 b of thelowest pressure turbine 19 of the turbines provided, and so correspondsto the most rearward turbine rotor in the direction of gas flow.

In the embodiment being described, the position of centre of gravity 108is measured as the axial distance between the intersection of theleading edge 64 a of one of the plurality of fan blades 64 and the hub66; and the centre of gravity of the gas turbine engine 10 as definedabove.

If a larger fan radius 102 is used, for example to improve propulsiveefficiency, such an increase may have an effect on the relative positionof the centre of gravity of the engine 10 should the engine componentssimply be scaled proportionally with the fan radius 102. This may causeproblems with mounting of the engine 10 to an aircraft wing 52 as theengine centre of gravity may be moved longitudinally away from the wing52. This may increase the load applied to a mounting pylon 53 connectingthe engine 10 and the wing 52.

In the embodiment shown in FIGS. 5A and 5B the centre of gravityposition ratio is in a range from 0.43 to 0.6. The skilled person willappreciate that the embodiments shown in FIGS. 5A and 5B are provided byway of examples falling within this range. More specifically, the centreof gravity position ratio may be in a range from 0.45 to 0.6 and morespecifically from 0.46 to 0.6. Yet more specifically, the centre ofgravity position ratio may be in a range from 0.47 to 0.49 or may be inrange from 0.45 to 0.48. The ranges in the previous sentence may, forexample, be for a gas turbine engine 10 with a fan tip radius 102 in therange from 110 cm to 150 cm or from 155 cm to 200 cm respectively.

The absolute values of the engine length 110 and centre of gravityposition 108 may be as defined elsewhere herein.

Defining the centre of gravity position ratio within the above rangesmay allow the centre of gravity to be located further rearwards comparedto the overall length of the engine 10. This may allow the centre ofgravity to be located at a position closer to a front mounting position53 a of the engine 10 (i.e. the position of a forward connection to apylon 53; in the embodiment being described the engine 10 is arranged tobe connected to a pylon 53 in two places, comprising a forward enginemount 53 a connecting the nacelle 21 to the pylon 53 and a rearwardengine mount 53 b connecting the core casing 11 a to the pylon 53. Theskilled person would appreciate that more, fewer, and/or differentmounting positions may be used in other embodiments). This may help toreduce or minimise mounting loads compared to centre of gravity positionratios found in known gas turbine engines or which would be achievedwith a proportional scaling of engine architecture. Other advantageouseffects such as reducing bending of the engine core 11 and deflection ofthe interconnecting shafts within the core may also be provided bydefining the centre of gravity position ratio as defined above.

By defining the centre of gravity position ratio within the rangedefined above the centre of gravity may be moved closer to a supportstructure (such as the pylon 53 of the embodiment being described)linking the engine core 11 and the nacelle 21. In the describedembodiment, the centre of gravity may be moved to a position in line (orclose to in line) to the fixed structure 24. This may reduce the forcetransmitted by the fixed structure 24 to support the engine core 11.

A fan speed to centre of gravity ratio of:

-   -   the centre of gravity position ratio×maximum take off rotational        fan speed        may be in a range from 600 rpm to 1350 rpm., and more        specifically from 650 rpm to 1276 rpm. For example for an engine        10 with a fan tip radius 102 in the range from 110 cm to 150 cm,        the fan speed to centre of gravity ratio may be 925 rpm to 1350        rpm. For an engine 10 with a fan tip radius 102 in the range        from 155 cm to 200 cm, the fan speed to centre of gravity ratio        may be 650 rpm to 910 rpm.

The maximum take-off rotational fan speed may be as defined elsewhereherein.

In the embodiment being described, the gas turbine engine 10 furthercomprises a gearbox 30 connected between the core shaft 26 and the fan23, the gearbox 30 being arranged to receive an input from the coreshaft 26 and providing an output to drive the fan at a lower rotationalspeed than the core shaft 26. In alternative embodiments, no gearbox maybe provided.

FIG. 5C illustrates a method 500 of operating an aircraft 70 comprisinga gas turbine engine 10 as described above.

The method comprises taking off 502, reaching cruise conditions 504, andcontrolling 506 the aircraft 70 such that the centre of gravity positionratio is in a range from 0.43 to 0.6, and using the engine to providethrust to the aircraft for take-off so that during take-off the fanspeed to centre of gravity ratio has a maximum value in a range asdescribed and/or claimed herein, for example from 600 rpm to 1350 rpm.

The centre of gravity position ratio and/or the fan speed to centre ofgravity ratio may more specifically be within any of the ranges definedabove (e.g. fan speed to centre of gravity ratio of 650 rpm to 1350rpm). The method 500 may include controlling the gas turbine engine 10according to any of the other parameters defined herein.

Gearbox Location to Engine Length Ratio

A gearbox location ratio may be defined as:

-   -   gearbox location (112)/engine length (110)

The engine length 110 may be measured as the axial distance between aforward region of the fan 23 and a rearward region of the lowestpressure turbine 19 (see FIGS. 6A and 6B).

In the embodiment being described, the engine length 110 is measured asthe axial distance between: the intersection of the leading edge 64 a ofone of the plurality of fan blades 64 and the hub 64; and a mean radiuspoint of the trailing edge of one of the rotor blades 44 of the lowestpressure turbine stage 19 b of the lowest pressure turbine 19 as definedabove. The mean radius point is the midpoint between a 0% span positionand a 100% span position of the rotor blade 44.

In the embodiment being described the gas turbine engine 10 has a singleturbine 19 referred to as the lowest pressure turbine. In otherembodiments, a plurality of turbines may be provided. The engine length110 is measured to a rotor of the lowest pressure stage 19 b of thelowest pressure turbine 19 of the turbines provided, and so correspondsto the most rearward turbine rotor in the direction of gas flow.

In the embodiment being described the gearbox location 112 is measuredas the axial distance between: the intersection of a leading edge 64 aof one of the fan blades 64 and the hub 66; and a radial planeintersecting the axial centre point of the ring gear 38 of the gearbox30 as defined above.

The gearbox 30 may contribute a large amount of the total mass of theengine 10. Its position along the length of the engine 10 may thereforehave a significant effect on the location of the centre of gravity.Should the components of the engine be scaled proportionally with anincreased fan size the relative position of the gearbox 30 may notprovide a suitable centre of gravity position 108 to allow efficientmounting of the engine 10 to an aircraft wing 52.

In the embodiment shown in FIGS. 6A and 6B the gearbox location ratio isin a range from 0.19 to 0.45. The skilled person will appreciate thatthe embodiments shown in FIGS. 6A and 6B are provided by way of examplesfalling within this range. In one embodiment, the gearbox location ratiomay be in a range from 0.19 to 0.3, and more specifically may be in arange from 0.19 to 0.25 or from 0.19 to 0.23. In one embodiment, thegearbox location ratio may be in a range from 0.19 to 0.23; this may be,for example, for an engine 10 with a fan tip radius 102 in the rangefrom 110 cm to 150 cm. In another embodiment, the gearbox location ratiomay be equal to or around 0.23; for example, in the range from 0.20 to0.25—this may be, for example, for an engine with a fan tip radius 102in the range from 155 cm to 200 cm.

The absolute values of the gearbox location 112 and engine length 110may be as defined elsewhere herein.

Defining the gearbox location ratio in the ranges above may allow orfacilitate control of the centre of gravity and assist engine mounting.A gearbox location ratio within the above range may cause an overallengine centre of gravity to be moved rearwards within the engine 10.This may allow the centre of gravity to be moved closer to the frontmounting position 53 a of the engine 10, and reduce front mounting loadscompared to known gas turbine engines 10 or which would be achieved witha proportional scaling of engine architecture. As already discussed,controlling the position of the centre of gravity in this way may alsoreduce bending of the engine core 11 and deflection of the core shaft26.

The choice of material from which the fan blades 64 are made may have animpact on the choice of gearbox location ratio. In the describedembodiment, the fan blades comprise a main body portion and a leadingedge portion. In embodiments where the main body portion of the fanblades 64 is formed at least partly from a composite material thegearbox location may be in a range between 50 cm to 110 cm and morespecifically in a range between 80 cm to 110 cm. The gearbox locationratio may be equal to or around 0.23 (e.g. in the range from 0.20 to0.25) where composite fan blades are used—this may be for an engine witha fan tip radius in the range from 155 cm to 200 cm.

In other embodiments, the fan blades 64 may be formed at least partlyfrom a metal or metal alloy. In one embodiment, the main body portion isformed from a metal alloy. The metal alloy may be, for example,aluminium-lithium alloy. In such embodiments the gearbox location may bein a range between 50 and 110 cm and more specifically may be in a rangebetween 50 cm and 80 cm. The gearbox location ratio may be in a rangefrom 0.19 to 0.23 where metallic fan blades are used. This may be, forexample, for an engine with a fan tip radius in the range from 110 cm to150 cm.

Outer Bypass to Fan Ratio

An outer bypass to fan ratio may be defined as:

$\frac{{the}\mspace{14mu}{outer}\mspace{14mu}{radius}\mspace{14mu}(114)\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{bypass}\mspace{14mu}{exhaust}\mspace{20mu}{nozzle}\mspace{14mu}(18)}{{the}\mspace{14mu}{fan}\mspace{14mu}{tip}\mspace{14mu}{radius}\mspace{14mu}(102)}$

In the embodiments being described, the outer bypass to fan ratio is inthe range from 0.6 to 1.05, and more particularly from 0.65 to 1.00. Invarious alternative embodiments, the outer bypass to fan ratio may belower than 1.05, optionally lower than 1.02, and further optionallylower than 1.00.

The fan tip radius 102 and the outer radius 114 of the bypass exhaustnozzle 18 are both as defined above—each radius is measured in a radialplane, perpendicular to the axis of the engine 10. In embodiments with afan tip radius 102 in the range from 110 cm to 150 cm, the outer bypassto fan ratio may be in the range from 0.95 to 1, and more particularly0.96 to 0.98. In embodiments with a fan tip radius 102 in the range from155 cm to 200 cm, the outer bypass to fan ratio may be in the range from0.91 to 0.98, optionally 0.94 to 0.96.

In the embodiment being described, the engine 10 comprises a nacelle 21and the fan tip radius 102 is approximately equal to the inner radius ofthe nacelle 21 adjacent the fan (in a forward region of the engine 10).The outer radius 114 of the bypass exhaust nozzle 18 is equivalent tothe inner radius of the nacelle 21 at the rearmost tip 21 b of thenacelle 21 (in a rearward region of the engine 10). The outer bypass tofan ratio therefore provides a measure of engine size variation fromfront to back.

In the embodiments being described, the bypass exhaust nozzle 18 has anexit plane 54 (marked in FIGS. 8A, 8B and 13A). The exit plane 54 is ina radial plane of the engine 10, perpendicular to the engine centreline9. The exit plane 54 extends inwardly from the rearmost tip of thenacelle 21. A flow area of the bypass exhaust nozzle 18 is approximatelydefined by the annular section of the exit plane 54 between the innersurface of the nacelle 21 and the outer surface of the engine core 11(i.e. the open part of the exit plane within the bypass duct 22/nozzle18, the nozzle 18 being the outlet of the duct 22, noting from thedefinitions above that the minimum distance R_(b) across the nozzle 18experienced by the bypass gas flow (B) may in fact be different from theradial nozzle width).

In the embodiments being described, the outer radius 114 of the bypassexhaust nozzle 18 is measured at the axial position of the exit plane 54of the bypass exhaust nozzle 18, which corresponds to the axial positionof the rearmost tip of the nacelle 21. The outer radius 114 of thebypass exhaust nozzle 18 is therefore the radial distance between thecentreline 9 of the engine 10 and an inner surface of the nacelle 21 atthe axial position of the rearmost tip 21 b of the nacelle 21.

In the embodiments being described, the outer radius 114 of the bypassexhaust nozzle 18 is approximately equal to, or smaller than, the fantip radius 102. In the embodiment shown in FIG. 7A, the outer radius 114of the bypass exhaust nozzle 18 is approximately equal to, but slightlylarger than, the fan tip radius 102, giving an outer bypass to fan ratioof 1.05 (Figures may not be to scale).

In the embodiment shown in FIG. 7B, the outer radius 114′ of the bypassexhaust nozzle 18 is smaller than the fan tip radius 102, giving anouter bypass to fan ratio of less than 1, more particularly between 0.9and 1, and more particularly of around 0.96 (figures may not be toscale).

The skilled person will appreciate that, in the embodiments shown inFIGS. 7A and 7B, the engine core 11 and fan 23 are identical, and thatthe difference in the outer bypass to fan ratio is due to the differentnacelle shape—and in particular to the inner surface of the nacelle 21curving inwards/towards the engine centre line towards the back of theengine 10 in the embodiment shown in FIG. 7B as opposed to curvingoutward/away from the engine centre line towards the back of the engine10 in the embodiment shown in FIG. 7A.

In the embodiment shown in FIG. 7A, an outer radius of the nacelle 21 isapproximately constant along the engine length 110, curving inwardslightly in the front and rear end regions only. By contrast, in theembodiment shown in FIG. 7B, the outer radius of the nacelle 21decreases from an axial mid-point of the nacelle 21 towards the rearportion. The nacelle 21 is also thinner than that of the embodimentshown in FIG. 7A, so providing a lower nacelle outer radius/diameter anda narrower overall engine 10 compared to size of the fan 23. The bypassexhaust duct 22 and exhaust nozzle 18 are therefore narrower in theembodiment shown in FIG. 7B.

The skilled person would appreciate that the relatively narrower andrearwardly inwardly-curving nacelle 21 may allow more room for a pylonstructure 53 connecting a rear portion of the engine 10 to an aircraftwing 52.

In alternative or additional embodiments, fan 23 parameters may bevaried to change the outer bypass to fan ratio, in addition to orinstead of changes to the nacelle 21.

Further, in various embodiments, parameters of the engine core 11 may bevaried so as to adjust bypass exhaust duct 22 and exhaust nozzle 18widths/flow areas independently of nacelle 21 radius (e.g. by making theinner radius 116 of the bypass exhaust nozzle 18 smaller).

FIG. 7C provides a schematic illustration of an engine 10 having anouter bypass to fan ratio in the range from 0.6 to 1.05. The skilledperson will appreciate that the embodiments shown in FIGS. 7A and 7B areprovided by way of examples falling within this range.

The skilled person would appreciate that having a relatively narrowbypass exhaust nozzle 18, as compared to fan size 102, may reduce dragproduced by the engine 10 in use. Further, the skilled person wouldappreciate that the relatively narrow bypass exhaust nozzle 18, and anoptionally correspondingly lower outer nacelle radius, may create a morecompact exhaust system, which may allow or facilitate under-winginstallation of a larger engine 10 on an aircraft 70.

In the embodiment being described, the gas turbine engine 10 furthercomprises a gearbox 30 connected between the core shaft 26 and the fan23, the gearbox 30 being arranged to receive an input from the coreshaft 26 and provide an output to drive the fan at a lower rotationalspeed than the core shaft 26. In the embodiment being described, thegearbox 30 has a gear ratio in the range of from 3 to 5, and moreparticularly of 3.2 to 3.8. In alternative embodiments, no gearbox maybe provided or the gear ratio may differ.

Inner Bypass to Fan Ratio

An inner bypass to fan ratio may be defined as:

$\frac{{the}\mspace{14mu}{inner}\mspace{14mu}{radius}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{bypass}\mspace{14mu}{exhaust}\mspace{20mu}{nozzle}\mspace{14mu}(18)}{{the}\mspace{14mu}{fan}\mspace{14mu}{tip}\mspace{14mu}{radius}\mspace{14mu}(102)}$

In the embodiments being described, the inner bypass to fan ratio is inthe range 0.4 to 0.65, and more particularly from 0.40 to 0.65. Inembodiments having an engine 10 with a fan tip radius 102 in the rangefrom 110 cm to 150 cm, the inner bypass to fan ratio may be in the rangefrom 0.57 to 0.63, for example being in the range from 0.58 to 0.60. Inembodiments having an engine 10 with a fan tip radius 102 in the rangefrom 155 cm to 200 cm, the inner bypass to fan ratio may be in the rangefrom 0.5 to 0.6, and optionally from 0.52 to 0.58.

The fan tip radius 102 and the inner radius 116 of the bypass exhaustnozzle 18 are both as defined above—each radius is measured in a radialplane, perpendicular to the axis 9 of the engine 10. The inner radius116 is measured in the same plane as the outer radius 114.

The fan tip radius 102 is approximately equal to the inner radius of thenacelle 21 adjacent the fan (in a forward region of the engine 10). Theinner radius 116 of the bypass exhaust nozzle 18 is equivalent to theouter radius of the engine core 11 at the axial position of the rearmosttip 21 b of the nacelle 21 (in a rearward region of the engine 10). Theinner bypass to fan ratio therefore provides a measure of engine sizevariation from front to back, differing from that of the outer bypass tofan ratio in that dimensions of the nacelle 21 are less important thanthose of the engine core 11.

In the embodiments being described, the bypass exhaust nozzle 18 has anexit plane 54 (marked in FIGS. 8A, 8B and 13A). The exit plane 54 is ina radial plane of the engine 10, perpendicular to the engine centreline9. The exit plane 54 extends inwardly from the rearmost tip of thenacelle 21. A flow area of the bypass exhaust nozzle 18 is approximatelydefined by the annular section of the exit plane 54 between the innersurface of the nacelle 21 and the outer surface of the engine core 11(i.e. the open part of the exit plane within the bypass duct 22/nozzle18, the nozzle 18 being the outlet of the duct 22, noting from thedefinitions above that the minimum distance R_(b) across the nozzle 18experienced by the bypass gas flow (B) may in fact be different from theradial nozzle width).

In the embodiments being described, the inner radius 116 of the bypassexhaust nozzle 18 is measured at the axial position of the exit plane 54of the bypass exhaust nozzle 18, which corresponds to the axial positionof the rearmost tip 21 b of the nacelle 21. The inner radius 116 of thebypass exhaust nozzle 18 is therefore the radial distance between thecentreline of the engine 10 and an outer surface of the engine core 11at the axial position of the rearmost tip of the nacelle 21/at the axialposition of the bypass exhaust nozzle exit plane 54.

In the embodiment being described, the inner radius 116 of the bypassexhaust nozzle 18 is in the range from 50 cm to 125 cm, and morespecifically from 65 cm to 110 cm. In embodiments having an engine 10with a fan tip radius 102 in the range from 110 cm to 150 cm, the innerradius of the bypass exhaust nozzle may be in the range from 65 cm to 90cm. In embodiments having an engine 10 with a fan tip radius 102 in therange from 155 cm to 200 cm, the inner radius of the bypass exhaustnozzle may be in the range from 80 cm to 110 cm.

In the embodiments being described, the inner radius 116 of the bypassexhaust nozzle 18 is smaller than the fan tip radius 102, for examplebeing around 50% of the fan tip radius. In the embodiment shown in FIG.8A, the inner radius 116 of the bypass exhaust nozzle 18 is over half ofthe length of the fan tip radius 102, giving an inner bypass to fanratio of around 0.6, and more particularly of around 0.64 (figures maynot be to scale). In the embodiment shown in FIG. 7B, the inner radius116′ of the bypass exhaust nozzle 18 is smaller than the fan tip radius102, giving an outer bypass to fan ratio of around 0.6, and moreparticularly of around 0.62 (figures may not be to scale).

The skilled person will appreciate that, in the embodiments shown inFIGS. 8A and 8B, the engine core 11 and fan 23 are identical, and thatthe difference in the inner bypass to fan ratio is due to the differentnacelle shape—and in particular to the nacelle 21 of the embodimentshown in FIG. 8B extending further back along the engine core 11 thanthat of FIG. 8A, so making the exit plane 54′ further back axially alongthe engine core 11. As the engine core radius decreases further backaxially along the engine core 11 in the embodiment shown, the innerradius 116 of the bypass exhaust nozzle 18 is smaller for the embodimentshown in FIG. 8B than for that in FIG. 8A. The skilled person willappreciate that inner radius of the nacelle 21 has no effect on themeasurement of the inner radius 116 of the bypass exhaust nozzle 18, butthat nacelle length does affect where the exit plane 54 of the bypassexhaust nozzle 18 is located, and therefore where the inner radius 116of the bypass exhaust nozzle 18 is measured. In alternative oradditional embodiments, shape of the engine core 11 may differ such thataxial position of the exit plane 54 has no effect, or a differenteffect, on the inner radius 116 of the bypass exhaust nozzle 18.

FIG. 8C provides a schematic illustration of an engine 10 having aninner bypass to fan ratio in the range from 0.4 to 0.65. The skilledperson will appreciate that the embodiments shown in FIGS. 8A and 8B areprovided by way of examples falling within this range.

The skilled person would appreciate that the engine core 11 is situatedradially within the bypass exhaust nozzle 18, and that the inner radius116 of the bypass exhaust nozzle 18 may therefore equivalently bethought of as an outer radius of the engine core 11. More generally,along the length of the engine core 11, the engine core 11 is situatedradially within the bypass exhaust duct 22 and an inner radius of thebypass exhaust duct 22 at any given axial location may thereforeequivalently be thought of as an outer radius of the engine core 11 atthat axial location.

The skilled person would appreciate that having a relatively narrowengine core 11, as compared to fan size 102, may reduce drag produced bythe engine 10 in use.

In the embodiment being described, the gas turbine engine 10 furthercomprises a gearbox 30 connected between the core shaft 26 and the fan23, the gearbox 30 being arranged to receive an input from the coreshaft 26 and provide an output to drive the fan at a lower rotationalspeed than the core shaft 26. In the embodiment being described, thegearbox 30 has a gear ratio in the range of from 3 to 5, and moreparticularly of 3.2 to 3.8. In alternative embodiments, no gearbox maybe provided or the gear ratio may differ.

Outer Bypass Duct Wall Angle Ratio

An outer bypass duct wall angle 126 is defined as described above. Inone embodiment, the outer bypass duct wall angle 126 may be in a rangebetween −15 to +1 degrees. The skilled person would appreciate thatFIGS. 12A, 12B and 12C are not to scale and are provided to show how theouter bypass duct wall angle is measured. By providing an outer BPD wallangle in this range a more compact exhaust system may be provided.

In one embodiment, the outer bypass duct wall angle may be negative. Inone embodiment, the outer bypass duct wall angle may be in a rangebetween −5 and −1 degrees. By using a negative angle so that the outerwall axis 60 slopes towards the engine centreline 9 a compact exhaustsystem may be provided. More specifically, the bypass duct wall anglemay be in a range between −4.0 to −1.0 degrees.

In one embodiment, the outer bypass duct wall angle 126 may be between−0.5 degrees and −4 degrees—this may be for an engine with a fan tipradius in the range from 110 cm to 150 cm. In one embodiment, the outerbypass duct wall angle may be in a range between −2.5 degrees and −4degrees—this may be for an engine with a fan tip radius 102 in the rangefrom 155 cm to 200 cm.

The radius of the bypass duct outlet guide vane (OGV) 58 used inembodiments having the bypass duct wall angle 126 defined within theranges above may be as defined elsewhere herein.

In the embodiment being described, the gas turbine engine 10 furthercomprises a gearbox 30 connected between the core shaft 26 and the fan23, the gearbox 30 being arranged to receive an input from the coreshaft 26 and providing an output to drive the fan at a lower rotationalspeed than the core shaft 26. In alternative embodiments, no gearbox maybe provided.

Fan Axis Angle

A fan axis angle 118 (also referred to as the fan tip axis angle) isdefined as described above. The fan axis angle 118 is defined by theangle between the fan tip axis 62 and the centreline 9 of the engine asshown in FIGS. 9A and 9B. The skilled person would appreciate that FIGS.9A and 9B are not to scale and are provided to show how the fan axisangle 118 is measured.

A positive value of the fan axis angle 118 corresponds to the fan tipaxis 62 sloping towards the engine centreline 9 when moving in arearward direction along the axis as illustrated in FIG. 9B, i.e. theradially outer tip 68 a of the leading edge 64 a the plurality of fanblades 64 is further from the engine centreline 9 compared to theradially outer tip of the trailing edge of the rotor blades 19 a of thelowest pressure stage of the turbine 19.

In the embodiment being described, the fan axis angle 118 is in a rangefrom 10 to 20 degrees. By providing a fan axis angle 118 in this rangethe gas turbine engine 10 may have a large fan diameter to provideimproved propulsive efficiency, whilst also having a relatively smalldiameter core 11.

In one embodiment, the fan axis angle may be in a range between 12degrees to 17 degrees. More specifically, the fan axis angle may be in arange between 13 degrees to 15 degrees. In one embodiment, the fan axisangle may be in a range between 13 degrees and 15 degrees—this may besuitable for an engine with a fan tip radius in the range from 110 cm to150 cm. In another embodiment, the fan axis angle may be in a rangebetween 13.5 degrees and 15.5 degrees—this may be suitable an enginewith a fan tip radius in the range from 155 cm to 200 cm.

In the embodiment being described the gas turbine engine 10 has a singleturbine 19 referred to as the lowest pressure turbine. In otherembodiments, a plurality of turbines may be provided. The fan axis angle118 is measured to a rotor of the lowest pressure stage 19 b of thelowest pressure turbine 19 of the turbines provided, and so correspondsto the most rearward turbine rotor 19 b in the direction of gas flow.

The values of the fan tip radius 102 and the turbine radius inembodiments having the fan axis angle 118 defined in the ranges abovemay be in the ranges defined elsewhere herein.

In the embodiment being described, the gas turbine engine 10 furthercomprises a gearbox 30 connected between the core shaft 26 and the fan23, the gearbox 30 being arranged to receive an input from the coreshaft 26 and providing an output to drive the fan at a lower rotationalspeed than the core shaft 26. In alternative embodiments, no gearbox maybe provided.

Fan Speed to Fan-Turbine Radius Difference Ratio

A fan speed to fan-turbine radius difference ratio is defined as:

$\frac{{the}\mspace{14mu}{maxiumum}\mspace{14mu}{take}\text{-}{off}\mspace{14mu}{rotational}\mspace{14mu}{speed}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{fan}}{{fan}\text{-}{turbine}\mspace{14mu}{radius}\mspace{14mu}{difference}\mspace{14mu}(120)}$

The maximum take-off rotational speed of the fan and the fan-turbineradius difference 120 are as defined above, and as illustrated in FIGS.10A and 10B.

In the embodiment being described in reference to FIGS. 10A and 10B, thefan speed to fan-turbine radius difference ratio is in a range between0.8 rpm/mm to 5 rpm/mm. As discussed above, this may reduce loading onthe pylon 53 which connects the engine 10 to the wing 52 of an aircraft70.

In one embodiment, the fan speed to fan-turbine radius ratio may be inrange between 1.5 rpm/mm to 4.0 rpm/mm. More specifically, the fan speedto fan-turbine radius ratio may be in range between 1.5 rpm/mm to 3.6rpm/mm. In one embodiment, the fan speed to fan-turbine radius ratio maybe in range between 2.93 rpm/mm and 3.8 rpm/mm—this may be for an engine10 with a fan tip radius 102 in the range from 110 cm to 150 cm. Inanother embodiment, the fan speed to fan-turbine radius ratio may be inrange between 1.2 rpm/mm and 2 rpm/mm—this may be for an engine 10 witha fan tip radius 102 in the range from 155 cm to 200 cm.

The fan-turbine radius difference 120 and the maximum take-offrotational speed of the fan 23 may be in the ranges defined elsewhereherein.

In the embodiment being described, with respect to the fan speed tofan-turbine radius difference ratio, the gas turbine engine 10 has asingle turbine 19 referred to as the lowest pressure turbine. In otherembodiments, a plurality of turbines may be provided. The fan speed tofan-turbine radius difference ratio is measured to a rotor of the lowestpressure stage 19 b of the lowest pressure turbine 19 of the turbinesprovided, and so corresponds to the most rearward turbine rotor in thedirection of gas flow.

The turbine radius 106 at the lowest pressure rotor stage, measured asthe radial distance from the engine centreline 9 to the radially outertip of the trailing edge of one of the rotor blades 44 of the lowestpressure stage 19 b of the turbine 19, may be in the range from 45 cm to85 cm. For an engine 10 with a fan tip radius 102 in a range from 110 cmto 150 cm, the turbine radius 106 at the lowest pressure rotor stage 19b may be in the range from 50 cm to 60 cm. For an engine 10 with a fantip radius 102 in a range from 155 cm to 200 cm, the turbine radius 106at the lowest pressure rotor stage 19 b may be in the range from rangefrom 60 cm to 85 cm.

In the embodiment being described, the gas turbine engine 10 furthercomprises a gearbox 30 connected between the core shaft 26 and the fan23, the gearbox 30 being arranged to receive an input from the coreshaft 26 and providing an output to drive the fan 23 at a lowerrotational speed than the core shaft 26. In alternative embodiments, nogearbox may be provided.

FIG. 10C illustrates a method 1000 of operating an aircraft 70comprising a gas turbine engine 10 as described above.

The method comprises taking off 1002, reaching cruise conditions 1004,and controlling 1006 the aircraft 70 such that the fan speed tofan-turbine radius ratio is in a range between 0.8 rpm/mm to 5 rpm/mmduring take-off. The fan speed to fan-turbine radius ratio may morespecifically be within any of the ranges defined above. The method mayinclude controlling the gas turbine engine 10 according to any of theother parameters defined herein.

Downstream Blockage Ratio

FIG. 11A provides a schematic illustration of an engine 10 located underthe wing 52 of an aircraft 70. The engine 10 is mounted to the wing by apylon 53. Any suitable pylon 53 known in the art may be used.

When on the ground, the aircraft 70 is arranged to rest on a groundplane 50. The skilled person would appreciate that lower surfaces oftyres of the aircraft's landing gear (not shown) generally make contactwith the ground plane 50. The wing 52 is arranged to lie a distance 124from the ground plane 50.

In the embodiment being described, the ground-to-wing distance 124 ismeasured between the ground plane 50 and the centre line of the wing 52at the leading edge 52 a of the wing 52.

The engine 10 is mounted beneath the wing 52, and positioned between thewing 52 and the ground 50 in normal operation. When the aircraft 70 ison the ground 50, the engine 10 is arranged to be below the wing 52 andabove the ground plane 50. The skilled person would appreciate that thediameter of the engine 10 is therefore arranged to be smaller than theground-to-wing distance 124 such that the engine 10 can be mountedbeneath the wing 52. The skilled person will appreciate that thediameter of the engine 10 is also arranged to allow space for a pylon 53to mount the engine to the wing.

In the embodiments being described, the engine 10 is arranged to extendforward of the leading edge 52 a of the wing 52. Only a rearward portionof the engine 10 therefore lies directly below the wing 52.

The turbine 19 lies below a forward region of the wing 52 in theembodiment being described, and more specifically below the leading edge52 a of the wing 52. In alternative embodiments, the turbine 19 may lieforward of, or rearward of, the leading edge 52 a of the wing 52. Thediameter 122 of the turbine 19, and more specifically the diameter 122of the turbine 19 at the axial position of the lowest pressure rotor 19b as defined above, therefore provides an indication of the amount ofvertical space below the wing 52 filled by the engine 10. The turbinediameter 122 is twice the turbine radius 106.

The amount of vertical space between the wing 52 and the ground plane 50taken up by the engine 10 may be described as a downstream blockage. Adownstream blockage ratio as defined above may therefore be calculatedas:

$\frac{\begin{matrix}{{the}\mspace{14mu}{turbine}\mspace{14mu}{diameter}\mspace{14mu}(122)\mspace{14mu}{at}\mspace{14mu}{an}\mspace{14mu}{axial}\mspace{14mu}{location}\mspace{14mu}{of}\mspace{14mu}{the}} \\{{lowest}\mspace{14mu}{pressure}\mspace{14mu}{rotor}\mspace{14mu}{stage}\mspace{14mu}\left( {19b} \right)}\end{matrix}}{{distance}\mspace{14mu}{between}\mspace{14mu}{ground}\mspace{14mu}{plane}\mspace{14mu}{and}\mspace{14mu}{wing}\mspace{14mu}(124)}$

In the embodiment being described, the downstream blockage ratio is inthe range from 0.2 to 0.3, more particularly in the range from 0.20 to0.30, in the range from 0.20 to 0.29 and particularly in the range from0.22 to 0.28. In embodiments with a fan tip radius 102 in the range from110 cm to 150 cm, the downstream blockage ratio may be in the range from0.23 to 0.25. In embodiments with a fan tip radius 102 in the range from155 cm to 200 cm, the downstream blockage ratio may be in the range from0.27 to 0.29.

In the embodiment being described, the turbine diameter 122 at the axiallocation of the lowest pressure rotor stage 19 b is as defined above andhas a value in one or more of the ranges defined above for turbinediameter.

In the embodiment being described, the turbine 19 is a first turbine 19,the compressor is a first compressor 14, and the core shaft is a firstcore shaft 26, and the engine core 11 further comprises a second turbine17, a second compressor 15, and a second core shaft 27 connecting thesecond turbine to the second compressor. In this embodiment, the secondturbine, second compressor, and second core shaft 27 are arranged torotate at a higher rotational speed than the first core shaft 26.

In the embodiment being described, the engine ratio of the engine 10 asdefined below falls within the ranges described below. In alternativeembodiments with a downstream blockage ratio in the range from 0.2 to0.3, the engine ratio may not fall within the range from 2.5 to 4—thefan diameter to engine length ratio may therefore not fall in the rangefrom 0.5 to 1.2.

Engine Ratio

An engine ratio may be defined as:

$\frac{\left( {{the}\mspace{14mu}{fan}\mspace{14mu}{{diameter}/{the}}\mspace{14mu}{engine}\mspace{14mu}{length}\mspace{14mu}(110)} \right)}{{the}\mspace{14mu}{downstream}\mspace{14mu}{blockage}\mspace{14mu}{ratio}} = \frac{\left( {2 \times {the}\mspace{14mu}{fan}\mspace{14mu}{radius}\mspace{14mu}{(102)/{the}}\mspace{14mu}{engine}\mspace{14mu}{length}\mspace{14mu}(110)} \right)}{{the}\mspace{14mu}{downstream}\mspace{14mu}{blockage}\mspace{14mu}{ratio}}$Where the engine length, fan radius and downstream blockage ratio areall as defined above.

In the embodiment being described, the engine ratio is in the range isin the range from 2.5 to 4, and more specifically in the range from 2.5to 4.0, and more specifically in the range from 2.7 to 3.7. In theembodiment being described, the engine ratio is greater than 2.5, andmore specifically greater than 3.0.

In the embodiment being described, the downstream blockage ratio of theengine 10 as defined above falls within the ranges described above. Inalternative embodiments with an engine ratio in the range from 2.5 to 4,the downstream blockage ratio may not fall in the range from 0.2 to0.3—i.e. the fan diameter to engine length ratio may not fall in therange from 0.5 to 1.2.

In the embodiment being described, the engine length 110 has a value inone or more of the absolute ranges defined above for engine length.

In the embodiment being described, the turbine diameter 122 at the axiallocation of the lowest pressure rotor stage 19 b has a value in one ormore of the absolute ranges defined above for turbine diameter.

In the embodiment being described, the turbine 19 is a first turbine 19,the compressor is a first compressor 14, and the core shaft is a firstcore shaft 26, and the engine core 11 further comprises a second turbine17, a second compressor 15, and a second core shaft 27 connecting thesecond turbine to the second compressor. In this embodiment, the secondturbine, second compressor, and second core shaft 27 are arranged torotate at a higher rotational speed than the first core shaft 26.

In the embodiment being described, a Q ratio defined as detailed belowis in a range from 0.005 Kgs⁻¹N⁻¹K^(1/2) to 0.011 Kgs⁻¹N⁻¹K^(1/2) andmore particularly from 0.006 Kgs⁻¹N⁻¹K^(1/2) to 0.009 Kgs⁻¹N⁻¹K^(1/2)where the value of Q is taken at cruise conditions.

Downstream Blockage and Q Ratio

In the embodiment illustrated in FIG. 11B, with reference to FIG. 14, aQ ratio of:

-   -   the downstream blockage ratio×Q        is in a range from 0.005 Kgs⁻¹N⁻¹K^(1/2) to 0.011        Kgs⁻¹N⁻¹K^(1/2), wherein the value of Q is taken at cruise        conditions.

The downstream blockage ratio and Q are as defined above. By definingthe Q ratio in this range a large mass flow may be achieved while alsominimising the downstream blockage. The Q ratio may also be representedas:

$\frac{\begin{matrix}{{the}\mspace{14mu}{turbine}\mspace{14mu}{diameter}\mspace{14mu}(122)\mspace{14mu}{at}\mspace{14mu}{an}\mspace{14mu}{axial}\mspace{14mu}{location}\mspace{14mu}{of}\mspace{14mu}{the}} \\{{lowest}\mspace{14mu}{pressure}\mspace{14mu}{rotor}\mspace{14mu}{stage}\mspace{14mu}\left( {19b} \right) \times Q}\end{matrix}}{{distance}\mspace{14mu}{between}\mspace{14mu}{ground}\mspace{14mu}{plane}\mspace{14mu}{and}\mspace{14mu}{wing}\mspace{14mu}(124)}$

In one embodiment, the Q ratio may be in a range from 0.005Kgs⁻¹N⁻¹K^(1/2) to 0.010 Kgs⁻¹N⁻¹K^(1/2). More specifically the Q ratiomay be in a range from 0.006 Kgs⁻¹N⁻¹K^(1/2) to 0.009 Kgs⁻¹N⁻¹K^(1/2).The Q value used in the ranges of the previous two sentences is taken atcruise conditions.

A specific thrust may be defined as net engine thrust divided by massflow rate through the engine. In one embodiment, at engine cruiseconditions:0.029Kgs ⁻¹ N ⁻¹ K ^(1/2) ≤Q≤0.036Kgs ⁻¹ N ⁻¹ K ^(1/2); and70Nkg ⁻¹ s≤specific thrust≤110Nkg ⁻¹ s.

In other embodiments, at cruise conditions: 0.032Kgs⁻¹N⁻¹K^(1/2)≤Q≤0.036 Kgs⁻¹N⁻¹K^(1/2). More specifically, at cruiseconditions: 0.033 Kgs⁻¹N⁻¹K^(1/2)≤Q≤0.035 Kgs⁻¹N⁻¹K^(1/2), or 0.034Kgs⁻¹N⁻¹K^(1/2)≤Q≤0.035 Kgs⁻¹N⁻¹K^(1/2)

The turbine diameter 122, ratio of the radius of fan blade at its hub tothe radius of the fan blade at its tip, and cruise conditions may be asdefined elsewhere herein.

In the embodiment being described, the turbine 19 is a first turbine 19,the compressor is a first compressor 14, and the core shaft is a firstcore shaft 26, and the engine core 11 further comprises a second turbine17, a second compressor 15, and a second core shaft 27 connecting thesecond turbine to the second compressor. In this embodiment, the secondturbine, second compressor, and second core shaft 27 are arranged torotate at a higher rotational speed than the first core shaft 26.

FIG. 14B illustrates a method 1400 of operating an aircraft 70comprising a gas turbine engine 10 as described above.

The method comprises taking off 1402, reaching cruise conditions 1404,and controlling 1406 the aircraft 70 such that the Q ratio is in a rangefrom 0.005 Kgs⁻¹N⁻¹K^(1/2) to 0.011 Kgs⁻¹N⁻¹K^(1/2) during take-off. TheQ ratio may more specifically be within any of the ranges defined above.The method may include controlling the gas turbine engine according toany of the other parameters defined herein.

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and sub-combinations of one or morefeatures described herein.

The invention claimed is:
 1. A gas turbine engine for an aircraft,comprises: an engine core comprising a turbine, a compressor, and a coreshaft connecting the turbine to the compressor, the engine core having acore length; a fan located upstream of the engine core, the fancomprising a plurality of fan blades extending radially from a hub, thehub and fan blades together defining a fan face having a fan face areaand a fan tip radius, a gearbox connected between the core shaft and thefan, the gearbox receiving an input from the core shaft and providing anoutput to drive the fan at a lower rotational speed than the core shaft,the turbine comprises a lowest pressure rotor stage, the turbine havinga turbine diameter at the lowest pressure rotor stage, wherein theturbine diameter at the lowest pressure rotor stage is measured at anaxial location of blade tip trailing edges of rotor blades of the lowestpressure rotor stage; and wherein an engine area ratio of:$\frac{{the}\mspace{14mu}{fan}\mspace{14mu}{face}\mspace{14mu}{area}}{{the}\mspace{14mu}{turbine}\mspace{14mu}{diameter} \times {the}\mspace{14mu}{core}\mspace{14mu}{length}}$is in a range from 1.7 to 3, and wherein, the gearbox has a gear ratioin a range of from 3.2 to
 5. 2. The gas turbine engine of claim 1,wherein the engine area ratio is in a range from 1.9 to 3.0.
 3. The gasturbine engine of claim 1, wherein the engine area ratio is in a rangefrom 2 to
 3. 4. The gas turbine engine of claim 1, wherein the enginearea ratio is in a range from 2.1 to 2.7.
 5. The gas turbine engine ofclaim 1, wherein the fan tip radius is measured between a centreline ofthe gas turbine engine and an outermost tip of each fan blade at itsleading edge, the fan face area being equal to π multiplied by thesquare of the fan tip radius.
 6. The gas turbine engine of claim 1,wherein the fan tip radius is in a range of from 110 cm to 200 cm. 7.The gas turbine engine of claim 1 wherein the fan tip radius is 110 cm,115 cm, 120 cm, 125 cm, 130 cm, 135 cm, 140 cm, 145 cm, 150 cm, 155 cm,160 cm, 165 cm, 170 cm, 175 cm, 180 cm, 185 cm, 190 cm or 195 cm.
 8. Thegas turbine engine of claim 1 wherein the fan tip radius is in a rangefrom 110 cm to 150 cm, and wherein the engine area ratio is in a rangefrom 1.7 to 2.7.
 9. The gas turbine engine of claim 1 wherein the fantip radius is in a range from 155 cm to 200 cm, and wherein the enginearea ratio is in a range from 2 to
 3. 10. The gas turbine engine ofclaim 1, wherein the turbine diameter of the turbine at the lowestpressure rotor stage is measured: (i) if rotor of the turbine isshrouded, to an underside of a shroud; or (ii) if the rotor isunshrouded, to the blade tips of the rotor.
 11. The gas turbine engineof claim 1, wherein the turbine diameter at the lowest pressure rotorstage is in a range from 70 cm to 170 cm.
 12. The gas turbine engine ofclaim 1, wherein a ratio of the fan tip radius to the turbine diameterat the lowest pressure rotor stage is in a range from 0.8 to 2.1. 13.The gas turbine engine of claim 1, wherein a ratio of the fan tip radiusto the core length is in a range from 0.3 to
 1. 14. The gas turbineengine of claim 1, wherein the core length is defined as an axialdistance between a forward region of the compressor and a rearwardregion of the turbine.
 15. The gas turbine engine of claim 1, whereinthe core length is measured along a centreline of the gas turbine enginefrom a mean radius point of a first stage of a compressor blade leadingedge to a mean radius point of the lowest pressure rotor stage bladetrailing edge of the turbine.
 16. The gas turbine engine of claim 1,wherein the engine core length is in a range from 150 cm to 320 cm. 17.The gas turbine engine of claim 1, wherein: each fan blade of theplurality of fan blades has a leading edge and a trailing edge; thelowest pressure turbine stage has a row of rotor blades, each of therotor blades extending radially and having a leading edge and a trailingedge; the gas turbine engine has a fan tip axis that joins a radiallyouter tip of the leading edge of one of the plurality of fan blades; andthe radially outer tip of the trailing edge of one or the rotor bladesof the lowest pressure stage of the turbine, the fan tip axis lying inlongitudinal plane which contains a centreline of the gas turbineengine, a fan tip axis angle is defined as the angle between the fan tipaxis and the centreline and the fan axis angle is in a range between 10and 20 degrees.
 18. The gas turbine engine according to claim 1,wherein: the turbine is a first turbine, the compressor is a firstcompressor, and the core shaft is a first core shaft; the engine corefurther comprises a second turbine, a second compressor, and a secondcore shaft connecting the second turbine to the second compressor; andthe second turbine, second compressor, and second core shaft arearranged to rotate at a higher rotational speed than the first coreshaft.